NOTICE OF
REINSTATEMENT
INCHPOUND
MIL-STD-1797A
NOTICE 3
24 August 2004
SUPERSEDING
NOTICE 2
19 December 1997
DEPARTMENT OF DEFENSE
INTERFACE STANDARD
FLYING QUALITIES OF PILOTED AIRCRAFT
MIL-STD-1797A, dated 30 January 1990, is hereby reinstated and may be used for acquisition.
Custodians:
Army – AV
Navy – AS
Air Force – 11
Preparing activity:
Air Force – 11
(Project 15GP-0066)
NOTE: The activities listed above were interested in this document as of the date of this document.
Since organizations and responsibilities can change, you should verify the currency of the information
above using the ASSIST Online database at www.dodssp.daps.mil.
AMSC N/A
DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.
FSC 15GP
NOTICE OF
CANCELLATION
INCH-POUND
MIL-STD-1797A
NOTICE 2
19 December 1997
DEPARTMENT OF DEFENSE
INTERFACE STANDARD
FLYING QUALITIES
OF
PILOTED AIRCRAFT
MIL-STD-1797A, dated 30 January 1990, is hereby canceled. However, the technical information and the
lessons learned from this standard have been preserved in MIL-HDBK-1797. MIL-HDBK-1797 is for
guidance only and should not be cited as a requirement, but information in the handbook may be useful in
determining or evaluating requirements.
Custodian:
Air Force - 11
Preparing Activity:
Air Force - 11
(Project 15GP-0043)
FSC 15GP
DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.
NOTICE OF
INCH-POUND
CHANGE
MIL-STD-1797A
NOTICE 1
28 June 1995
DEPARTMENT OF DEFENSE
INTERFACE STANDARD
FLYING QUALITIES
OF
PILOTED AIRCRAFT
TO ALL HOLDERS OF MIL-STD-1797A:
MAKE THE FOLLOWING PEN AND INK CHANGES:
Cover
Delete “Military Standard” and insert “Department of Defense Interface Standard.” Delete entire
Distribution Statement D and substitute: “Distribution Statement A. Approved for public release;
distribution is unlimited.”
Page ii
Delete entire Export Control Warning and Destruction Notice.
Page 9
3.4.6, bottom of page. Delete “ ∆β ” and substitute “ ∆βmax ”.
Page 38
4.6.5.3, line 3, change to read: “system due to pilot action shall not exceed…”.
Page 39
4.6.6.2, first sentence, change to read: “Yaw control power shall be sufficient…”.
Page 58
20.3, sixth reference on page. Delete “MIL-F-25140” and substitute “MIL-W-25140”.
DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.
AMSC N/A
FSC 15GP
i
MIL-STD-1797A
NOTICE 1
Page 80
3.2, Requirement Guidance, Category B, line 3. Delete "light-path" and substitute "flight-path".
Page 96
4.1.4.1, Requirement Guidance, d. Delete "VominPA" and substitute “ Vomin PA ”.
4.1.4.1, Requirement Guidance, d, lines 2, 3, 4. and 6. Delete "Vs "and substitute “VS”.
Page 101
4.1.4.2, Requirement Guidance, 2.e, first line. Delete "what" and substitute “that".
Page 103
4.1.4.3, Requirement Guidance, next to last paragraph, line 3. Change to read: "…sea level while
remaining…”.
4.1.4.3, Requirement Guidance, next to last paragraph, line 7. Change to read: “…(commonly used for
structural…”.
Page 122
Figure 8, Change title to read: "Definition of Levels which include atmospheric disturbances as well as
failures -Suggested by Carlson (AFFDL-TR-78-171)."
Page 137
5.1.11.2, Verification Guidance, last sentence. Change to read: “… will likely be at the boundaries of...”.
Page 139
4.1.11.4, Requirement Guidance, line 2. Change to read: "Recommended minimum time delay: A default
value would be 1 second.”
4.1.11.4, Requirement Guidance, last sentence. Change to read: “Table IX and the paragraph following it
are excerpts.'
Page 144
4.1.11.5. Requirement Guidance, third paragraph, line 2. Delete " C1β ” and substitute “Clβ”.
Page 146
5.1.11.5, Verification Guidance, third paragraph, lines 2 and 3. Change to read: “ …To these margins
must be added another nose-down control…”.
5.1.11.5, Verification Guidance, last paragraph, line 2. Change to read: "needed: ∆δmarg is the sum of
turbulence and sensor-wise components, ∆δtran provides…”.
Page 159
4.1.12. 11 line 3. Change to read: “...aircraft motion produced, be conveniently and ...”.
2
MIL-STD-1797A
NOTICE 1
Page 160
4.1.12.4 through 4.1.12.11, Requirement Rationale, first line. Change to read: “…directly to the flight
control system applies…”.
Page 167
4.2.1.1, Requirement Guidance, Equation, last term in numerator. Change to read " e
−τ
es
".
Page 169
4.2 through 4.2.1.1, Requirement Lessons Learned, third paragraph, line 1. Change to read "... angle of
attack above that for zero 1 / Th1 .or dγ / dV . These…".
Page 175
4.2.1.2, Requirement Guidance, second and third expressions. Change to read:
(
)
Mα [ζ p ; ωp ]
Mδ Z w 1 Tθ1
Phugoid
and
( ) −τθs
s [ζ sp ; ωsp ]
Mδ 1 Tθ2 e
Short Period
Page 177
4.2.1.2, Requirement Guidance, line 8. Change to read "...can be ignored, leaving the az/δes
numerator…”.
4.2.1.2, Requirement Guidance, last paragraph, line 4. Change to read “…(LOES) of the θ/Fes transfer
function…”.
Page 184
4.2.1.2, Requirement Guidance, last line. Change to read “…short-period approximation),".
Page 187
4.2.1.2, Requirement Guidance, line 2. Change to read “…and 16 ftc, the specified Level 1…”.
Page 194
Figure 25, approximately center of figure. Delete “15,2" and substitute "1.5,2".
Figure 25, legend box, P.R. Scale, second listing. Delete "COPPER-HARPER" and substitute "COOPERHARPER”.
Figure 25, title. Change to read: "Comparison of pilot ratings with Category A short-period frequency
requirements."
Page 195
Figure 26, legend box, P.R. Scale, second listing. Delete "COPPER-HARPER" and substitute "COOPERHARPER".
3
MIL-STD-1797A
NOTICE 1
Page 196
Figure 27, legend box, P.R. Scale, second listing. Delete "COPPER-HARPER" and substitute “COOPERHARPER".
Page 197
Figure 29, legend box, P.R. Scale, second listing. Delete "COPPER-HARPER" and substitute ”COOPERHARPER".
Figure 29, right edge of figure. Delete "Boundaries from AGARD-CP-333" and substitute "Boundaries
from AGARD-CP-333, Gibson".
Page 200
Figures 33 and 35, fourth diagonal line. Delete “LEVELS 2" and substitute "LEVEL 2".
Figure 35, legend in upper left hand corner, last line. Delete "4,000 43" and substitute "4,000 431".
Page 201
Figure 36. Rotate the figure 90 degrees counterclockwise.
Page 219
Next to last equation. Delete and substitute:
ω 2sp
n/α
=
&θ&
0
n z ss
≅
q ss ∆t
g
−
q ss VT g
V T ∆t
Page 259
Frequency-response magnitude and phase, second equation under Exact: Delete and substitute:
Φ = ± tan −1 (τω)
Page 260
Equation at bottom of page. Delete and substitute:
ζ = K q TqM / 2
δ
Page 290
4.2.7.2.1, last sentence. Change to read: “…result in departure or exceedance of load factor units."
Page 294
4.2.7.3, Lessons Learned, pages 2 and 3. Change to read: "…lt was possible to mis-set trim for takeoff so
that -24 deg deflection ...”.
4
MIL-STD-1797A
NOTICE 1
Page 303
Table XVII, last line of section a. Change to read: "* For nL <3, (Fs / n) max is 28.0...”.
Page 305
Figure 101, center of figure, denominator, twice. Delete "(n/d)" and substitute "(n/α)".
Page 329
Table XV111, last column heading. Delete "das/Fas" and substitute “δas/Fas”.
Page 364
Figure 130, legend in lower right comer. Delete "REFERENCE AFFDL-TR-65-227" and "REFERENCE
NASA-TND-2251"and substitute "AFFDL-TR-65-227" and "NASA-TND-2251”.
Page 366
5.3.1.2, Verification Guidance, second line of equation. Delete “-sin (a+ it)" and substitute "-sin (α+ it)".
Page 375
4.4.1.1, Requirement Guidance, fifth paragraph, line 6. Change expression to read: “(M2cos2∆ - 1)-1/2”.
Page 377
4.5.1.1, Requirement Guidance, first equation following second paragraph. Change denominator in
second part of equation to read: “(1/TS) (1/TR) [ζ d , ω d ] ”.
4.5.1.1, Requirement Guidance, second equation following second paragraph. Change numerator in
second part of equation to read: “ A 3 s 3 + A 2 s 2 + A 1s + A o e-τeβ”.
(
)
Page 391
First equation, last term, right hand side. Change to read: “ − 2k 2z C L 1 ”.
Page 398
Figure 150, title. Change to read: "Composite pilot ratings for spiral descent of simulated reentry vehicle
(from NASA-CR-778).”
Page 402
5.5.1.3, Verification Guidance, two large equations in middle of page. Change the denominator inside the
large parentheses of both equations to read:
1−
gρbk 2z
4( W
/ S)k 2x
C′

 C y + nr
 β 2k 2
z

 C ′l p

 C ′n
 β
Page 404
4.5.1.4, Requirement Guidance. third paragraph, second line. Delete “figure 154” and substitute “figure
156".
5
MIL-STD-1797A
NOTICE 1
Page 418
Second equation. last term inside the square brackets on the right side.
Delete −
C y ′δ
C′l δ
a
a
C′l β
and substitute
−
Cy δ
C′l δ
a
C′l β
a
Page 429
5.5.5, Verification Guidance, last equation, denominator. Change to read:
 Cn C l
δa
δr
Cl δ 1 −
a 
C l δ Cn δ
a
r





Page 460
5.5.8.1, Verification Guidance, last two equations, bottom of page. Change to read:
(
)
(
)
k 2x = I x / mb 2 , k 2z = I z / mb 2 ,
Page 464
Lower right hand comer, top of second sketch. Delete “X” and substitute “X".
Page 466
4.5.8.4, Requirement Lessons Learned, line 2. Delete “YF-15" and substitute “YC-15".
4.5.8.4, Requirement Lessons Learned. Combine the two paragraphs into one paragraph.
Page 476
4.5.9.2, Requirement Guidance, fifth line following table. Change to read “Level 2: One-eighth of the Level
2 values in table XXXVI.”
Page 491
4.5.9.3, Requirement Guidance, fifth line from bottom of the page. Delete “(AGARD-C-333)" and
substitute “(AGARD-CP-333)”.
Page 498
Matrix equation, middle of page. Change to read:
6
MIL-STD-1797A
NOTICE 1
2 


 C lr
 2
bg
2 c 

C l β +
C
4
k
k
sin Φ tan Φ  C l δ
−
−

yβ
z
y
2
2


a

2 VT
b 

 C L1



 Cn
bg

C nβ +
C y β  r − 4k 2xz sin Φ tan Φ 
C nδ

2
a
2 VT

 C L1








β
 =
δ a 

 1

c2 
− Cl r sin Φ + CL1  k 2z − k 2y 2  sin2 Φ tan Φ 
bg  2
b 



VT2 
1
− Cnr sin Φ + CL1k 2xz sin2 Φ tan Φ


2


Sentence following the matrix equation. Change to read:
(
)
(
)
(
)
where C L 1 = w / (q S ), k 2y = I y / m c 2 , k 2z = I z / mb 2 , k 2x z = I xz / mb 2 ; n y = C yβ β / C L 1 .
Page 504
4.5.9.5.7, Requirement Guidance, third paragraph, line 2. Delete "over look” and substitute "overlook”.
Page 505
4.6.1.1, Requirement Guidance, second and third equations, first term in denominator. Change to read:
“(s + 1/TS)
4.6.1.1, Requirement Guidance, first line following third equation. Change to read ".. deflection controls
(pilot controller deflection commands ...”.
Page 526
5.6.1.1, Verification Guidance, fourth equation. Change to read:
2ζ d ωd =& −
k 2 C′l β
1
ρVg 
C yβ + 2 C′nr − 2z
2( W / S) 
2k z
k x C′n β


 CL − 1 C′n 
1
p
2


2k z


5.6.1.1, Verification Guidance, first and second sentences following last equation. Change to
read φ / β d is the ratio of amplitudes of the roll and sideslip envelopes in the dutch roll mode
and C L 1 = w / (q S ) . The dutch-roll envelope of roll rate, p, is shown in figure 228, from AFFDLTR-69-72, for a step command."
Page 529
Last equation, numerator of last term. Change to read:
7
(
)
C y δ C l β − Cn β C l δ / Cn δ .
a
r
r
MIL-STD-1797A
NOTICE 1
Page 545
4.6.2, Requirement Guidance, footnote 11, second sentence. Change to read: "The ψβ and ∆β/k of figure
238…”.
Page 546
Paragraph 3, line 3. Change to read: “… For φ / β d above some nominal value (=& 5.0 ) ,…”.
Page 552
Second line following figure. Change to read: “measurement of φ1. The above-noted trend…”.
Page 572
5.6.4, Verification Guidance, first equation. Change to read: − δ a δ rp = C′l δ
r
C′l δ
a
Page 578
5.6.5.1, Verification Guidance, last three equations. Change to read:
C *yβ = C yβ − C y δ C hrβ C hrδ
r
C n* β = C nβ − C nδ C hrβ C hrδ
r
C l* β = C l β − C l δ C hrβ C hrδ
r
r
r
r
Page 581
5.6.6, Verification Guidance. equation. Change to read:
q SbCn(δr, δa, δT, β,
β& , P, R, α, M …) = IZ R& - Ixz ( P& + QR) + (Iy - Ix) PQ
Page 584
4.6.6.2, first sentence. Change to read: "Yaw control power shall be sufficient to meet…”.
Page 589
First equation, delete and substitute.
C
 yβ
C
 lβ
C
 nβ
0
Cl δ
a
Cn δ
a
Cyδ 
r
Cl δ 
r 
Cn δ 
r


C y p sin θ / cos φ − C y r cos θ
β


  g b
 δa  = ⋅
⋅ sin φ ⋅  Cl p sin θ / cos φ − Cl r cos φ + 2C L1 cos θ sin φ tan φ k 2z − k 2y c 2 / b2 cos θ + k 2xz sin θ / cos φ 


  V 2V
C sin θ / cos φ − C cos θ + 2C cos θ sin φ tan φ k 2 − k 2 c 2 / b2 sin θ / cos φ + k 2 cos θ 
 δr 
nr
L1
x
y
xz

 n p
[(
[(
8
)
)
]
]
MIL-STD-1797A
NOTICE 1
Second equation, delete and substitute.
CL1 = W /( q S) ,
k 2x = I x /(mb 2 ),
k 2y = I y /(m c 2 ),
k 2z = I z /(mb 2 ),
k 2xz = I xz /(mb 2 )
First paragraph, line 4. Change to read "hand column (0 C2 C3)T (neglecting C yp and C y r ),”.
Fourth equation, delete and substitute.
δa =
[C (C
gb sin Φ
2
2V 2 ∆
y β C nδr
)
(
− C nβ C y δ − C 3 C yβ C l δ − C l β C y δ
r
r
r
)]
Page 603
Equation inside sketch. Delete "z = x tan ε” and substitute “-z = x tan ε”.
4.8. 1, Requirement Lessons Learned . Switch the order of paragraphs 1 and 2.
Page 604
2
5.8.1, Verification Guidance, last sentence. Change to read: “…the critical roll rate squared, p cr
, is".
Page 642
5.8.4.3.1, Verification Guidance, four equations, one-third of the page down from the top. Change to read:
Li +
L ′i =
I xz
Ni
Ix
I2
1 − xz
IxIz
Ni +
, N′i =
I xz
Li
Iz
I2
1 − xz
I xIz
, C ′l i =
Cli +
I xz
C ni
Iz
I2
1 − xz
I xIz
, C ′ni =
C ni +
1−
I xz
C li
Ix
I 2xz
IxIz
Page 664
Figure 266, top of figure. Close the gap in the line under the “h”.
Page 665
Fourth and fifth equations. Delete and substitute:
φu1 (Ω) =
φv1 (Ω) =
200
(ft/sec)2 per rad/ft
1 + (100Ω )2
939 [1 + ( 400Ω2 )]
2
2
[1 + (1000Ω ) ] [1 + (400Ω / 3) ]
(ft/sec)2 per rad/ft
Page 667
First and second sentence following third equation. Delete and substitute:
where: ωp = Ship pitch frequency, radians/second.
θs = Ship pitch amplitude, radians.
9
MIL-STD-1797A
NOTICE 1
Page 674
Table LVI, second column, fourth entry, line 4. Change to read: “components. High-frequency
spectral”.
Page 680
4.9.3, first equation. Delete and substitute:
- α& g = qg =
∂wg
∂wg
∂vg
, pg =
, rg = −
∂x
∂y
∂x
Page 693
First row, fourth column. Change to read "iii. 1,687".
Second row, fourth column. Change to read "687”.
Page 701
Column heading, fourth column. Change to read “PAGE NO. IN THIS DOCUMENT”.
10
MIL-STD-1797A
NOTICE 1
2. THE FOLLOWING PAGES OF MIL-STD-1797A HAVE BEEN REVISED AND SUPERSEDE
THE PAGES LISTED:
NEW PAGE
DATE
SUPERSEDED PAGE
DATE
iii
30 January 1990
iii
REPRINTED WITHOUT CHANGE
iv-xii
28 June 1995
iv-xxvii
30 January 1990
61
28 June 1995
61
REPRINTED WITHOUT CHANGE
62-75b
28 June 1995
62-75
30 January 1990
76
30 January 1990
76
REPRINTED WITHOUT CHANGE
107
30 January 1990
107
REPRINTED WITHOUT CHANGE
108-108q
28 June 1995
108
30 January 1990
109
28 June 1995
109
30 January 1990
110
30 January 1990
110
REPRINTED WITHOUT CHANGE
149
28 June 1995
149
30 January 1990
150
30 January 1990
150
REPRINTED WITHOUT CHANGE
151
28 June 1995
151
30 January 1990
152-152d
28 June 1995
152
30 January 1990
267-275
28 June 1995
267-275
30 January 1990
276
30 January 1990
276
REPRINTED WITHOUT CHANGE
281
28 June 1995
281
30 January 1990
282
28 June 1995
282
30 January 1990
421
30 January 1990
421
REPRINTED WITHOUT CHANGE
422-422a
28 June 1995
422
30 January 1990
569
30 January 1990
569
REPRINTED WITHOUT CHANGE
570
28 June 1995
570
30 January 1990
3. RETAIN THIS NOTICE AND INSERT BEFORE TABLE OF CONTENTS.
4. Holders of MIL-STD-1797A will verity that corrections, page changes, and additions indicated above
have been entered. This notice page will be retained as a check sheet. This issuance, together with
appended pages, is a separate publication. Each notice is to be retained by stocking points until the
military standard is completely revised or canceled.
Custodian:
Preparing Activity:
Army - AV
Air Force - 11
Navy - AS
Air Force - 11
(Project 15GP-0111)
11
MIL-STD-1797A
FOREWORD
This standard is intended for use with fixed-wing aircraft supported primarily by aerodynamic force rather
than engine thrust. It also covers the handling characteristics of aircraft under piloted control on the
ground, and may be used with powered-lift aircraft in aerodynamic flight (above the conversion speed,
Vcon). This standard also applies to piloted transatmospheric flight when flight depends upon aerodynamic
lift and/or air breathing propulsion systems. Flying qualities of military rotorcraft are specified in MIL-H8501. while flying qualities in V/STOL flight are the subject of MIL-F-83300.
For further background information, see Appendix C.
REPRINTED WITHOUT CHANGE
iii
MIL-STD-1797A
CONTENTS
Paragraph
Number
Title
Standard
Page
1.
1.1
1.2
SCOPE ...................................................................................................................................... 1
Purpose. .................................................................................................................................... 1
Applicability................................................................................................................................ 1
2.
2.1
2.1.1
2.1.2
2.2
2.3
APPLICABLE DOCUMENTS ................................................................................................... 1
Government documents. ........................................................................................................... 1
Specifications, standards, and handbooks................................................................................ 1
Other Government documents, drawings, and publications. .................................................... 1
Non-Government publications. .................................................................................................. 1
Order of precedence.................................................................................................................. 1
3.
3.1
3.2
3.3
3.4
3.4.1
3.4.2
3.4.3
3.4.4
3.4.5
3.4.6
3.4.7
3.5
DEFINITIONS ............................................................................................................................ 2
Aircraft classification and operational missions......................................................................... 2
Flight Phase Categories. ........................................................................................................... 2
Levels and qualitative suitability of flying qualities. ................................................................... 2
Parameters. ............................................................................................................................... 2
General terms............................................................................................................................ 2
Speeds....................................................................................................................................... 3
Thrust and power....................................................................................................................... 5
Control parameters.................................................................................................................... 6
Longitudinal parameters ............................................................................................................ 6
Lateral-directional parameters................................................................................................... 8
Atmospheric disturbance parameters...................................................................................... 16
Terms used in high angle of attack requirements ................................................................... 18
4.
4.1
4.1.1
4.1.2
4.1.3
4.1.4
4.1.4.1
4.1.4.2
4.1.4.3
4.1.5
4.1.6
4.1.6.1
4.1.6.2
4.1.6.3
4.1.7
4.1.7.1
4.1.7.2
4.1.7.3
4.1.7.4
4.1.7.5
4.1.7.6
4.1.8
4.1.8.1
4.1.8.2
4.1.9
4.1.10
4.1.11
REQUIREMENTS.................................................................................................................... 19
General requirements.............................................................................................................. 19
Loadings. ................................................................................................................................. 19
Moments and products of inertia. ............................................................................................ 19
Internal and external stores. .................................................................................................... 19
Flight Envelopes ...................................................................................................................... 19
Operational Flight Envelopes. ................................................................................................. 19
Service Flight Envelopes. ........................................................................................................ 19
Permissible Flight Envelopes. ................................................................................................. 19
Configurations and States of the aircraft. ................................................................................ 19
Aircraft Normal States. ............................................................................................................ 22
Allowable Levels for Aircraft Normal States. ........................................................................... 22
Flight outside the Service Flight Envelopes. ........................................................................... 22
Ground operation..................................................................................................................... 22
Aircraft Failure States. ............................................................................................................. 22
Allowable Levels for Aircraft Failure States............................................................................. 22
Aircraft Special Failure States. ................................................................................................ 22
Probability calculation.............................................................................................................. 22
Generic failure analysis. .......................................................................................................... 25
When Levels are not specified. ............................................................................................... 25
Failures outside the Service Flight Envelopes. ....................................................................... 25
Dangerous flight conditions. .................................................................................................... 25
Warning and indication. ........................................................................................................... 25
Devices for indication, warning, prevention, and recovery...................................................... 26
Interpretation of subjective requirements. ............................................................................... 26
Interpretation of quantitative requirements.............................................................................. 26
General flying qualities requirements ...................................................................................... 26
Supersedes pages iv - xiii of
MIL-STD-1797A
iv
MIL-STD-1797A
CONTENTS
Paragraph
Number
4.1.11.1
4.1.11.2
4.1.11.3
4.1.11.4
4.1.11.5
4.1.11.6
4.1.11.7
4.1.11.8
4.1.12
4.1.12.1
4.1.12.2
4.1.12.3
4.1.12.4
4.1.12.5
4.1.12.6
4.1.12.7
4.1.12.8
4.1.12.9
4.1.12.10
4.1.12.11
4.1.13
4.1.13.1
4.1.13.2
4.1.13.3
4.1.13.4
4.1.13.5
4.1.13.6
4.2
4.2.1
4.2.1.1
4.2.1.2
4.2.2
4.2.3
4.2.4
4.2.5
4.2.6
4.2.6.1
4.2.6.2
4.2.7
4.2.7.1
4.2.7.2
4.2.7.2.1
4.2.7.3
4.2.7.4
4.2.8
4.2.8.1
4.2.8.2
4.2.8.3
4.2.8.4
4.2.8.5
4.2.8.6
4.2.8.6.1
Title
Standard
Page
Buffet. ...................................................................................................................................... 26
Release of stores..................................................................................................................... 26
Effects of armament delivery and special equipment.............................................................. 26
Failures. ................................................................................................................................... 26
Control margin. ........................................................................................................................ 27
Pilot-induced oscillations (PIO)................................................................................................ 27
Residual oscillations. ............................................................................................................... 27
Control cross-coupling............................................................................................................. 27
General flight control system characteristics........................................................................... 27
Control centering and breakout forces. ................................................................................... 27
Cockpit control free play. ......................................................................................................... 28
Adjustable controls. ................................................................................................................. 28
Rate of control displacement. .................................................................................................. 28
Dynamic characteristics........................................................................................................... 28
Damping. ................................................................................................................................. 28
Transfer to alternate control modes. ....................................................................................... 28
Flight control system failures. .................................................................................................. 28
Augmentation systems. ........................................................................................................... 28
Auxiliary dive recovery devices. .............................................................................................. 29
Direct force controllers............................................................................................................. 29
General trim requirements....................................................................................................... 29
Trim system irreversibility. ....................................................................................................... 29
Rate of trim operation. ............................................................................................................. 29
Stalling of trim systems............................................................................................................ 29
Transients and trim changes. .................................................................................................. 29
Trim for asymmetric thrust....................................................................................................... 30
Automatic trim system. ............................................................................................................ 30
Flying qualities requirements for the pitch axis. ...................................................................... 30
Pitch attitude dynamic response to pitch controller ................................................................. 30
Long-term pitch response........................................................................................................ 30
Short-term pitch response. ...................................................................................................... 30
Pilot-induced pitch oscillations................................................................................................. 30
Residual pitch oscillations. ...................................................................................................... 30
Normal acceleration at pilot station. ........................................................................................ 30
Pitch trim changes. .................................................................................................................. 30
Pitch axis response to other inputs ......................................................................................... 31
Pitch axis response to failures, controls free........................................................................... 31
Pitch axis response to configuration or control mode change................................................. 31
Pitch axis control power........................................................................................................... 33
Pitch axis control power in unaccelerated flight. ..................................................................... 33
Pitch axis control power in maneuvering flight. ....................................................................... 33
Load factor response............................................................................................................... 33
Pitch axis control power in takeoff........................................................................................... 33
Pitch axis control power in landing. ......................................................................................... 33
Pitch axis control forces........................................................................................................... 33
Pitch axis control forces--steady-state control force per g. ..................................................... 33
Pitch axis control forces--transient control force per g. ........................................................... 33
Pitch axis control forces--control force variations during rapid speed changes. .................... 34
Pitch axis control forces--control force vs. control deflection. ................................................. 34
Pitch axis control breakout forces. .......................................................................................... 34
Pitch axis control force limits ................................................................................................... 34
Pitch axis control force limits--takeoff...................................................................................... 34
Supersedes pages iv - xiii of
MIL-STD-1797A
v
MIL-STD-1797A
CONTENTS
Paragraph
Number
4.2.8.6.2
4.2.8.6.3
4.2.8.6.4
4.2.8.6.5
4.2.8.6.6
4.2.8.7
4.2.9
4.2.9.1
4.2.9.2
4.3
4.3.1
4.3.1.1
4.3.1.2
4.3.2
4.3.3
4.3.3.1
4.3.3.2
4.3.4
4.4
4.4.1
4.4.1.1
4.5
4.5.1
4.5.1.1
4.5.1.2
4.5.1.3
4.5.1.4
4.5.1.5
4.5.2
4.5.3
4.5.4
4.5.5
4.5.6
4.5.7
4.5.7.1
4.5.7.2
4.5.8
4.5.8.1
4.5.8.2
4.5.8.3
4.5.8.4
4.5.8.5
4.5.8.6
4.5.9
4.5.9.1
4.5.9.2
4.5.9.3
4.5.9.4
4.5.9.5
4.5.9.5.1
4.5.9.5.2
4.5.9.5.3
Title
Standard
Page
Pitch axis control force limits--landing. .................................................................................... 34
Pitch axis control force limits--dives. ....................................................................................... 34
Pitch axis control force limits-sideslips. ................................................................................... 34
Pitch axis control force limits-failures. ..................................................................................... 35
Pitch axis control force limits-control mode change. ............................................................... 35
Pitch axis trim systems. ........................................................................................................... 35
Pitch axis control displacements ............................................................................................. 35
Pitch axis control displacements-takeoff. ................................................................................ 35
Pitch axis control displacements-maneuvering. ...................................................................... 35
Flying qualities requirements for the normal (flight path) axis................................................. 35
Flight path response to attitude change .................................................................................. 35
Transient flight path response to attitude change. .................................................................. 35
Steady-state flight path response to attitude change. ............................................................. 35
Flight path response to designated flight path controller......................................................... 36
Flight path control power ......................................................................................................... 36
Control power for designated primary flight path controller..................................................... 36
Control power for designated secondary flight path controller. ............................................... 36
Flight path controller characteristics........................................................................................ 36
Flying qualities requirements for the longitudinal (speed) axis ............................................... 36
Speed response to attitude changes....................................................................................... 36
Speed response to attitude changes-relaxation in transonic flight.......................................... 36
Flying qualities requirements for the roll axis .......................................................................... 36
Roll response to roll controller ................................................................................................. 36
Roll mode................................................................................................................................. 36
Spiral stability........................................................................................................................... 36
Coupled roll-spiral oscillation................................................................................................... 36
Roll oscillations........................................................................................................................ 37
Roll time delay. ........................................................................................................................ 37
Pilot-induced roll oscillations. .................................................................................................. 37
Linearity of roll response to roll controller................................................................................ 37
Lateral acceleration at the pilot station.................................................................................... 37
Roll characteristics in steady sideslip...................................................................................... 37
Roll axis control for takeoff and landing in crosswinds............................................................ 37
Roll axis response to other inputs ........................................................................................... 37
Roll axis response to augmentation failures............................................................................ 37
Roll axis response to configuration or control mode change. ................................................. 37
Roll axis control power ............................................................................................................ 38
Roll axis response to roll control inputs................................................................................... 38
Roll axis control power in steady sideslips. ............................................................................. 38
Roll axis control power in crosswinds...................................................................................... 38
Roll axis control power for asymmetric thrust.......................................................................... 38
Roll axis control power in dives and pullouts........................................................................... 38
Roll axis control power for asymmetric loading. ...................................................................... 38
Roll axis control forces and displacements ............................................................................. 39
Roll control displacements....................................................................................................... 39
Roll axis control forces to achieve required roll performance. ................................................ 39
Roll axis control sensitivity....................................................................................................... 39
Roll axis control centering and breakout forces. ..................................................................... 39
Roll axis control force limits ..................................................................................................... 39
Roll axis control force limits in steady turns. ........................................................................... 39
Roll axis control force limits in dives and pullouts. .................................................................. 39
Roll axis control force limits in crosswinds. ............................................................................. 39
Supersedes pages iv - xiii of
MIL-STD-1797A
vi
MIL-STD-1797A
CONTENTS
Paragraph
Number
4.5.9.5.4
4.5.9.5.5
4.5.9.5.6
4.5.9.5.7
4.6
4.6.1
4.6.1.1
4.6.1.2
4.6.1.3
4.6.2
4.6.3
4.6.4
4.6.5
4.6.5.1
4.6.5.2
4.6.5.3
4.6.6
4.6.6.1
4.6.6.2
4.6.6.3
4.6.7
4.6.7.1
4.6.7.2
4.6.7.3
4.6.7.4
4.6.7.5
4.6.7.6
4.6.7.7
4.6.7.8
4.6.7.9
4.6.7.10
4.6.7.11
4.7
4.7.1
4.8
4.8.1
4.8.2
4.8.3
4.8.4
4.8.4.1
4.8.4.2
4.8.4.2.1
4.8.4.2.2
4.8.4.2.3
4.8.4.2.4
4.8.4.3
4.8.4.3.1
4.8.4.3.2
4.9
4.9.1
4.9.2
4.9.3
Title
Standard
Page
Roll axis control force limits in steady sideslips....................................................................... 39
Roll axis control force limits for asymmetric thrust. ................................................................. 39
Roll axis control force limits for failures. .................................................................................. 39
Roll axis control force limits for configuration or control mode change................................... 40
Flying qualities requirements for the yaw axis......................................................................... 40
Yaw axis response to yaw and side-force controllers ............................................................. 40
Dynamic lateral-directional response. ..................................................................................... 40
Steady sideslips....................................................................................................................... 40
Wings-level turn....................................................................................................................... 40
Yaw axis response to roll controller......................................................................................... 41
Pilot-induced yaw oscillations.................................................................................................. 41
Yaw axis control for takeoff and landing in crosswinds........................................................... 41
Yaw axis response to other inputs .......................................................................................... 41
Yaw axis response to asymmetric thrust................................................................................. 41
Yaw axis response to failures.................................................................................................. 42
Yaw axis response to configuration or control mode change.................................................. 42
Yaw axis control power............................................................................................................ 42
Yaw axis control power for takeoff, landing, and taxi. ............................................................. 42
Yaw axis control power for asymmetric thrust......................................................................... 43
Yaw axis control power with asymmetric loading. ................................................................... 43
Yaw axis control forces............................................................................................................ 43
Yaw axis control force limits in rolling maneuvers................................................................... 43
Yaw axis control force limits in steady turns............................................................................ 43
Yaw axis control force limits during speed changes. .............................................................. 43
Yaw axis control force limits in crosswinds. ............................................................................ 43
Yaw axis control force limits with asymmetric loading. ........................................................... 44
Yaw axis control force limits in dives and pullouts. ................................................................. 44
Yaw axis control force limits for waveoff (go-around). ............................................................ 44
Yaw axis control force limits for asymmetric thrust during takeoff. ......................................... 44
Yaw axis control force limits with flight control failures. .......................................................... 44
Yaw axis control force limits-control mode change. ................................................................ 44
Yaw axis breakout forces. ....................................................................................................... 44
Flying qualities requirements for the lateral flight path axis .................................................... 44
Dynamic response for lateral translation. ................................................................................ 44
Flying qualities requirements for combined axes .................................................................... 45
Cross-axis coupling in roll maneuvers..................................................................................... 45
Crosstalk between pitch and roll controllers............................................................................ 45
Control harmony. ..................................................................................................................... 45
Flight at high angle of attack.................................................................................................... 45
Warning cues........................................................................................................................... 45
Stalls. ....................................................................................................................................... 45
Stall approach.......................................................................................................................... 45
Stall characteristics.................................................................................................................. 45
Stall prevention and recovery. ................................................................................................. 46
One-engine-out stalls. ............................................................................................................. 46
Post-stall gyrations and spins.................................................................................................. 46
Departure from controlled flight. .............................................................................................. 46
Recovery from post-stall gyrations and spins.......................................................................... 46
Flying qualities requirements in atmospheric disturbances..................................................... 47
Allowable flying qualities degradations in atmospheric disturbances. .................................... 47
Definition of atmospheric disturbance model form. ................................................................. 48
Application of disturbance models in analyses........................................................................ 48
Supersedes pages iv - xiii of
MIL-STD-1797A
vii
MIL-STD-1797A
CONTENTS
Paragraph
Number
5.
5.1
5.1.1
5.1.2
5.1.3
5.1.4
5.1.4.1
5.1.4.2
5.1.4.3
5.1.5
5.1.6
5.1.6.1
5.1.6.2
5.1.6.3
5.1.7
5.1.7.1
5.1.7.2
5.1.7.3
5.1.7.4
5.1.7.5
5.1.7.6
5.1.8
5.1.8.1
5.1.8.2
5.1.9
5.1.10
5.1.11
5.1.11.1
5.1.11.2
5.1.11.3
5.1.11.4
5.1.11.5
5.1.11.6
5.1.11.7
5.1.11.8
5.1.12
5.1.12.1
5.1.12.2
5.1.12.3
5.1.12.4
5.1.12.5
5.1.12.6
5.1.12.7
5.1.12.8
5.1.12.9
5.1.12.10
5.1.12.11
5.1.13
5.1.13.1
5.1.13.2
5.1.13.3
Title
Standard
Page
VERIFICATION ....................................................................................................................... 49
General requirements-verification ........................................................................................... 49
Loadings-verification................................................................................................................ 49
Moments and products of inertia-verification........................................................................... 49
Internal and external stores-verification. ................................................................................. 49
Flight Envelopes-verification.................................................................................................... 49
Operational Flight Envelopes-verification................................................................................ 49
Service Flight Envelopes-verification. ..................................................................................... 49
Permissible Flight Envelopes-verification................................................................................ 49
Configurations and States of the aircraft-verification. ............................................................. 49
Aircraft Normal States-verification........................................................................................... 49
Allowable Levels for Aircraft Normal States-verification.......................................................... 49
Flight outside the Service Flight Envelopes-verification.......................................................... 49
Ground operation-verification. ................................................................................................. 49
Aircraft Failure States-verification. .......................................................................................... 49
Allowable Levels for Aircraft Failure States-verification .......................................................... 50
Aircraft Special Failure States-verification............................................................................... 50
Probability calculation-verification. .......................................................................................... 50
Generic failure analysis-verification......................................................................................... 50
When Levels are not specified-verification.............................................................................. 50
Failures outside the Service Flight Envelopes-verification...................................................... 50
Dangerous flight conditions-verification................................................................................... 50
Warning and indication-verification. ........................................................................................ 50
Devices for indication, warning, prevention, and recovery-verification. .................................. 50
Interpretation of subjective requirements-verification.............................................................. 50
Interpretation of quantitative requirements-verification. .......................................................... 50
General flying qualities requirements-verification.................................................................... 50
Buffet-verification..................................................................................................................... 50
Release of stores-verification. ................................................................................................. 50
Effects or armament delivery and special equipment-verification. .......................................... 50
Failures-verification. ................................................................................................................ 50
Control margin-verification....................................................................................................... 50
Pilot-induced oscillations(PIO)-verification. ............................................................................. 51
Residual oscillations-verification. ............................................................................................ 51
Control cross-coupling-verification. ......................................................................................... 51
General flight control system characteristics-verification. ....................................................... 51
Control centering and breakout forces-verification.................................................................. 51
Cockpit control free play-verification. ...................................................................................... 51
Adjustable controls-verification................................................................................................ 51
Rate of control displacement-verification. ............................................................................... 51
Dynamic characteristics-verification. ....................................................................................... 51
Damping-verification................................................................................................................ 51
Transfer to alternate control modes-verification. ..................................................................... 51
Flight control system failures-verification. ............................................................................... 51
Augmentation systems-verification.......................................................................................... 51
Auxiliary dive recovery devices-verification............................................................................. 51
Direct force controllers-verification. ......................................................................................... 51
General trim requirements-verification .................................................................................... 51
Trim system irreversibility-verification. .................................................................................... 51
Rate of trim operation-verification. .......................................................................................... 52
Stalling of trim systems-verification. ........................................................................................ 52
Supersedes pages iv - xiii of
MIL-STD-1797A
viii
MIL-STD-1797A
CONTENTS
Paragraph
Number
5.1.13.4
5.1.13.5
5.1.13.6
5.2
5.2.1
5.2.1.1
5.2.1.2
5.2.2
5.2.3
5.2.4
5.2.5
5.2.6
5.2.6.1
5.2.6.2
5.2.7
5.2.7.1
5.2.7.2
5.2.7.2.1
5.2.7.3
5.2.7.4
5.2.8
5.2.8.1
5.2.8.2
5.2.8.3
5.2.8.4
5.2.8.5
5.2.8.6
5.2.8.6.1
5.2.8.7
5.2.8.7.1
5.2.8.7.2
5.2.8.7.3
5.2.8.7.4
5.2.8.8
5.2.9
5.2.9.1
5.2.9.2
5.3
5.3.1
5.3.1.1
5.3.1.2
5.3.2
5.3.3
5.3.3.1
5.3.3.2
5.3.4
5.4
5.4.1
5.4.1.1
5.5
5.5.1
Title
Standard
Page
Transients and trim changes-verification................................................................................. 52
Trim for asymmetric thrust-verification. ................................................................................... 52
Automatic trim system-verification........................................................................................... 52
Flying qualities requirements for the pitch axis-verification..................................................... 52
Pitch attitude dynamic response to pitch controller-verification .............................................. 52
Long-term pitch response-verification. .................................................................................... 52
Short-term pitch response-verification..................................................................................... 52
Pilot-induced pitch oscillations-verification. ............................................................................. 52
Residual pitch oscillations-verification..................................................................................... 52
Normal acceleration at pilot station-verification....................................................................... 52
Pitch trim changes-verification................................................................................................. 52
Pitch axis response to other inputs-verification ....................................................................... 52
Pitch axis response to failures, controls free-verification. ....................................................... 52
Pitch axis response to configuration or control mode change-verification. ............................. 52
Pitch axis control power-verification ........................................................................................ 52
Pitch axis control power in unaccelerated flight-verification.................................................... 52
Pitch axis control power in maneuvering flight-verification...................................................... 52
Load factor response-verification. ........................................................................................... 53
Pitch axis control power in takeoff-verification. ....................................................................... 53
Pitch axis control power in landing-verification........................................................................ 53
Pitch axis control forces-verification ........................................................................................ 53
Pitch axis control forces-steady-state control force per g-verification. .................................... 53
Pitch axis control forces-transient control force per g-verification........................................... 53
Pitch axis control forces-control force variations during rapid speed changes
-verification. ............................................................................................................................. 53
Pitch axis control forces-control force vs. control deflection-verification................................. 53
Pitch axis control breakout forces-verification. ........................................................................ 53
Pitch axis control force limits-verification................................................................................. 53
Pitch axis control force limits-takeoff-verification..................................................................... 53
Pitch axis control force limits-landing-verification.................................................................... 53
Pitch axis control force limits-dives-verification. ...................................................................... 53
Pitch axis control force limits-sideslips-verification.................................................................. 53
Pitch axis control force limits-failures-verification.................................................................... 53
Pitch axis control force limits-control mode change-verification.............................................. 53
Pitch axis trim systems-verification. ........................................................................................ 53
Pitch axis control displacements-verification........................................................................... 54
Pitch axis control displacements-takeoff-verification............................................................... 54
Pitch axis control displacements-maneuvering-verification..................................................... 54
Flying qualities requirements for the normal (flight path) axis-verification .............................. 54
Flight path response to attitude change-verification................................................................ 54
Transient flight path response to attitude change-verification................................................. 54
Steady-state flight path response to attitude change-verification............................................ 54
Flight path response to designated flight path controller-verification. ..................................... 54
Flight path control power-verification....................................................................................... 54
Control power for designated primary flight path controller-verification. ................................. 54
Control power for designated secondary flight path controller-verification. ............................ 54
Flight path controller characteristics-verification. .................................................................... 54
Flying qualities requirements for the longitudinal (speed) axis-verification............................. 54
Speed response to attitude changes-verification. ................................................................... 54
Speed response to attitude changes-relaxation in transonic flight-verification. ...................... 54
Flying qualities requirements for the roll axis-verification........................................................ 54
Roll response to roll controller-verification .............................................................................. 54
Supersedes pages iv - xiii of
MIL-STD-1797A
ix
MIL-STD-1797A
CONTENTS
Paragraph
Number
5.5.1.1
5.5.1.2
5.5.1.3
5.5.1.4
5.5.1.5
5.5.2
5.5.3
5.5.4
5.5.5
5.5.6
5.5.7
5.5.7.1
5.5.7.2
5.5.8
5.5.8.1
5.5.8.2
5.5.8.3
5.5.8.4
5.5.8.5
5.5.8.6
5.5.9
5.5.9.1
5.5.9.2
5.5.9.3
5.5.9.4
5.5.9.5
5.5.9.5.1
5.5.9.5.2
5.5.9.5.3
5.5.9.5.4
5.5.9.5.5
5.5.9.5.6
5.5.9.5.7
5.6
5.6.1
5.6.1.1
5.6.1.2
5.6.1.3
5.6.2
5.6.3
5.6.4
5.6.5
5.6.5.1
5.6.5.2
5.6.5.3
5.6.6
5.6.6.1
5.6.6.2
5.6.6.3
5.6.7
5.6.7.1
5.6.7.2
Title
Standard
Page
Roll mode-verification. ............................................................................................................. 54
Spiral stability-verification. ....................................................................................................... 54
Coupled roll-spiral oscillation-verification. ............................................................................... 55
Roll oscillations-verification. .................................................................................................... 55
Roll time delay-verification....................................................................................................... 55
Pilot-induced roll oscillations-verification................................................................................. 55
Linearity of roll response to roll controller-verification. ............................................................ 55
Lateral acceleration at the pilot station-verification. ................................................................ 55
Roll characteristics in steady sideslip-verification. .................................................................. 55
Roll axis control for takeoff and landing in crosswinds-verification. ........................................ 55
Roll axis response to other inputs-verification......................................................................... 55
Roll axis response to augmentation failures-verification. ........................................................ 55
Roll axis response to configuration or control mode change-verification................................ 55
Roll axis control power-verification.......................................................................................... 55
Roll axis response to roll control inputs-verification. ............................................................... 55
Roll axis control power in steady sideslips-verification. .......................................................... 55
Roll axis control power in crosswinds-verification. .................................................................. 55
Roll axis control power for asymmetric thrust-verification. ...................................................... 55
Roll axis control power in dives and pullouts-verification. ....................................................... 55
Roll axis control power for asymmetric loading-verification. ................................................... 55
Roll axis control forces and displacements-verification........................................................... 56
Roll control displacements-verification. ................................................................................... 56
Roll axis control forces to achieve required roll performance-verification............................... 56
Roll axis control sensitivity-verification. ................................................................................... 56
Roll axis control centering and breakout forces-verification.................................................... 56
Roll axis control force limits-verification .................................................................................. 56
Roll axis control force limits in steady turns-verification.......................................................... 56
Roll axis control force limits in dives and pullouts-verification................................................. 56
Roll axis control force limits in crosswinds-verification............................................................ 56
Roll axis control force limits in steady sideslips-verification. ................................................... 56
Roll axis control force limits for asymmetric thrust-verification................................................ 56
Roll axis control force limits for failures-verification................................................................. 56
Roll axis control force limits for configuration or control mode change-verification. ............... 56
Flying qualities requirements for the yaw axis-verification ...................................................... 56
Yaw axis response to yaw and side-force controllers-verification........................................... 56
Dynamic lateral-directional response-verification.................................................................... 56
Steady sideslips-verification. ................................................................................................... 56
Wings-level turn-verification. ................................................................................................... 56
Yaw axis response to roll controller-verification. ..................................................................... 56
Pilot-induced yaw oscillations-verification. .............................................................................. 57
Yaw axis control for takeoff and landing in crosswinds-verification. ....................................... 57
Yaw axis response to other inputs-verification ........................................................................ 57
Yaw axis response to asymmetric thrust-verification. ............................................................. 57
Yaw axis response to failures-verification. .............................................................................. 57
Yaw axis response to configuration or control mode change-verification. .............................. 57
Yaw axis control power-verification. ........................................................................................ 57
Yaw axis control power for takeoff, landing, and taxi-verification............................................ 57
Yaw axis control power for asymmetric thrust-verification. ..................................................... 57
Yaw axis control power with asymmetric loading-verification. ................................................ 57
Yaw axis control forces-verification. ........................................................................................ 57
Yaw axis control force limits in rolling maneuvers-verification. ............................................... 57
Yaw axis control force limits in steady turns-verification. ........................................................ 57
Supersedes pages iv - xiii of
MIL-STD-1797A
x
MIL-STD-1797A
CONTENTS
Paragraph
Number
Title
Standard
Page
5.6.7.3
5.6.7.4
5.6.7.5
5.6.7.6
5.6.7.7
5.6.7.8
5.6.7.9
5.6.7.10
5.6.7.11
5.7
5.7.1
5.8
5.8.1
5.8.2
5.8.3
5.8.4
5.8.4.1
5.8.4.2
5.8.4.2.1
5.8.4.2.2
5.8.4.2.3
5.8.4.2.4
5.8.4.3
5.8.4.3.1
5.8.4.3.2
5.9
5.9.1
5.9.2
5.9.3
Yaw axis control force limits during speed changes-verification. ............................................ 57
Yaw axis control force limits in crosswinds-verification........................................................... 57
Yaw axis control force limits with asymmetric loading-verification. ......................................... 57
Yaw axis control force limits in dives and pullouts-verification................................................ 57
Yaw axis control force limits for waveoff (go-around)-verification. .......................................... 57
Yaw axis control force limits for asymmetric thrust during takeoff-verification........................ 58
Yaw axis control force limits with flight control failures-verification. ........................................ 58
Yaw axis control force limits-control mode change-verification............................................... 58
Yaw axis breakout forces-verification...................................................................................... 58
Flying qualities requirements for the lateral flight path axis-verification .................................. 58
Dynamic response for lateral translation-verification. ............................................................. 58
Flying qualities requirements for combined axes-verification.................................................. 58
Cross-axis coupling in roll maneuvers-verification. ................................................................. 58
Crosstalk between pitch and roll controllers-verification. ........................................................ 58
Control harmony-verification.................................................................................................... 58
Flight at high angle of attack-verification ................................................................................. 58
Warning cues-verification. ....................................................................................................... 58
Stalls-verification. .................................................................................................................... 58
Stall approach-verification. ...................................................................................................... 58
Stall characteristics-verification. .............................................................................................. 58
Stall prevention and recovery-verification. .............................................................................. 58
One-engine-out stalls-verification............................................................................................ 58
Post-stall gyrations and spins-verification. .............................................................................. 58
Departure from controlled flight-verification............................................................................. 58
Recovery from post-stall gyrations and spins-verification. ...................................................... 59
Flying qualities requirements in atmospheric disturbances-verification .................................. 59
Allowable flying qualities degradations in atmospheric disturbances-verification. .................. 59
Definition of atmospheric disturbance model form-verification................................................ 59
Application of disturbance models in analyses-verification. .................................................... 59
6.
6.1
6.2
6.3
6.4
6.5
6.6
6.7
NOTES .................................................................................................................................... 59
Intended use............................................................................................................................ 59
Level definitions....................................................................................................................... 59
Reference documents tree. ..................................................................................................... 59
Data requirements. .................................................................................................................. 60
Subject term (key word) listing ................................................................................................ 61
Responsible engineering office (REO). ................................................................................... 61
Changes from previous issue. ................................................................................................. 61
Supersedes pages iv - xiii of
MIL-STD-1797A
xi
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
Title
Page
10.
10.1
10.2
10.3
10.4
10.4.1
10.4.2
10.5
SCOPE .................................................................................................................................... 69
Scope....................................................................................................................................... 69
Purpose. .................................................................................................................................. 69
Use. ......................................................................................................................................... 69
Format ..................................................................................................................................... 69
Requirement/verification identity. ............................................................................................ 69
Requirement/verification package. .......................................................................................... 69
Responsible engineering office. .............................................................................................. 69
20.
20.1
20.2
20.3
APPLICABLE DOCUMENTS ................................................................................................. 69
References. ............................................................................................................................. 69
Avoidance of tiering. ................................................................................................................ 69
Government documents .......................................................................................................... 69
30.
3.1
3.2
3.3
DEFINITIONS .......................................................................................................................... 88
Aircraft classification and operational missions....................................................................... 88
Flight Phase Categories. ......................................................................................................... 92
Levels and qualitative suitability of flying qualities. ................................................................. 97
40.
4.1
4.1.1
4.1.2
4.1.3
4.1.4
4.1.4.2
4.1.4.3
4.1.5
4.1.6
4.1.6.1
4.1.6.2
4.1.6.3
4.1.7
4.1.7.1
4.1.7.2
4.1.7.3
4.1.7.4
4.1.7.5
4.1.7.6
4.1.8
4.1.8.1
4.1.8.2
4.1.9
4.1.10
4.1.11
4.1.11.1
4.1.11.2
4.1.11.3
4.1.11.4
4.1.11.5
4.1.11.6
4.1.11.7
REQUIREMENTS.................................................................................................................. 101
General requirements............................................................................................................ 101
Loadings. ............................................................................................................................... 101
Moments and products of inertia. .......................................................................................... 103
Internal and external stores ................................................................................................... 105
Flight Envelopes. ................................................................................................................... 106
Service Flight Envelopes. ...................................................................................................... 112
Permissible Flight Envelopes. ............................................................................................... 115
Configurations and States of the aircraft. .............................................................................. 117
Aircraft Normal States. .......................................................................................................... 119
Allowable Levels for Aircraft Normal States. ......................................................................... 120
Flight outside the Service Flight Envelopes. ......................................................................... 120
Ground operation................................................................................................................... 120
Aircraft Failure States. ........................................................................................................... 122
Allowable Levels for Aircraft Failure States........................................................................... 123
Aircraft Special Failure States. .............................................................................................. 124
Probability calculation............................................................................................................ 126
Generic failure analysis. ........................................................................................................ 135
When Levels are not specified. ............................................................................................. 138
Failures outside the Service Flight Envelopes. ..................................................................... 139
Dangerous flight conditions. .................................................................................................. 140
Warning and indication. ......................................................................................................... 140
Devices for indication, warning, prevention, and recovery.................................................... 140
Interpretation of subjective requirements. ............................................................................. 142
Interpretation of quantitative requirements............................................................................ 143
General flying qualities requirements .................................................................................... 146
Buffet. .................................................................................................................................... 146
Release of stores................................................................................................................... 149
Effects of armament delivery and special equipment............................................................ 150
Failures. ................................................................................................................................. 151
Control margin. ...................................................................................................................... 155
Pilot-induced oscillations (PIO).............................................................................................. 163
Residual oscillations. ............................................................................................................. 165
Supersedes pages iv - xiii of
MIL-STD-1797A
xii
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
4.1.11.8
4.1.12
4.1.12.1
4.1.12.2
4.1.12.3
4.1.12.4
4.1.12.5
4.1.12.6
4.1.12.7
4.1.12.8
4.1.12.9
4.1.12.10
4.1.12.11
4.1.13
4.1.13.1
4.1.13.2
4.1.13.3
4.1.13.4
4.1.13.5
4.1.13.6
4.2
4.2.1
4.2.1.1
4.2.1.2
4.2.2
4.2.3
4.2.4
4.2.5
4.2.6.
4.2.6.1
4.2.6.2
4.2.7
4.2.7.1
4.2.7.2
4.2.7.2.1
4.2.7.3
4.2.7.4
4.2.8
4.2.8.1
4.2.8.2
4.2.8.3
4.2.8.4
4.2.8.5
4.2.8.6
4.2.8.6.1
4.2.8.6.2
4.2.8.6.3
4.2.8.6.4
4.2.8.6.5
4.2.8.6.6
4.2.8.7
4.2.9
Title
Page
Control cross-coupling........................................................................................................... 166
General flight control system characteristics......................................................................... 168
Control centering and breakout forces. ................................................................................. 168
Cockpit control free play. ....................................................................................................... 168
Adjustable controls. ............................................................................................................... 168
Rate of control displacement. ................................................................................................ 171
Dynamic characteristics......................................................................................................... 171
Damping. ............................................................................................................................... 171
Transfer to alternate control modes. ..................................................................................... 171
Flight control system failures. ................................................................................................ 171
Augmentation systems. ......................................................................................................... 171
Auxiliary dive recovery devices. ............................................................................................ 171
Direct force controllers.......................................................................................................... 171
General trim requirements..................................................................................................... 176
Trim system irreversibility. ..................................................................................................... 176
Rate of trim operation. ........................................................................................................... 176
Stalling of trim systems.......................................................................................................... 176
Transients and trim changes. ................................................................................................ 176
Trim for asymmetric thrust..................................................................................................... 176
Automatic trim system. .......................................................................................................... 176
Flying qualities requirements for the pitch axis ..................................................................... 179
Pitch attitude dynamic response to pitch controller ............................................................... 179
Long-term pitch response...................................................................................................... 179
Short-term pitch response. .................................................................................................... 183
Pilot-induced pitch oscillations............................................................................................... 281
Residual pitch oscillations. .................................................................................................... 290
Normal acceleration at pilot station. ...................................................................................... 292
Pitch trim changes. ................................................................................................................ 295
Pitch axis response to other inputs. ...................................................................................... 297
Pitch axis response to failures. controls free........................................................................ 297
Pitch axis response to configuration or control mode change............................................... 299
Pitch axis control power......................................................................................................... 301
Pitch axis control power in unaccelerated flight. ................................................................... 301
Pitch axis control power in maneuvering flight. ..................................................................... 302
Load factor response............................................................................................................. 304
Pitch axis control power in takeoff......................................................................................... 306
Pitch axis control power in landing. ....................................................................................... 309
Pitch axis control forces......................................................................................................... 311
Pitch axis control forces--steady-state control force per g. ................................................... 316
Pitch axis control forces--transient control force per g. ......................................................... 332
Pitch axis control forces--control force variations during rapid speed changes. ................... 340
Pitch axis control forces--control force vs. control deflection. ............................................... 341
Pitch axis control breakout forces. ........................................................................................ 350
Pitch axis control force limits ................................................................................................. 351
Pitch axis control force limits -- takeoff.................................................................................. 351
Pitch axis control force limits--landing. .................................................................................. 353
Pitch axis control force limits--dives. ..................................................................................... 355
Pitch axis control force limits--sideslips................................................................................. 357
Pitch axis control force limits--failures. .................................................................................. 359
Pitch axis control force limits--control mode change............................................................. 360
Pitch axis trim systems. ......................................................................................................... 361
Pitch axis control displacements ........................................................................................... 363
Supersedes pages iv - xiii of
MIL-STD-1797A
xiii
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
4.2.9.1
4.2.9.2
4.3
4.3.1
4.3.1.1
4.3.1.2
4.3.2
4.3.3
4.3.3.1
4.3.3.2
4.3.4
4.4
4.4.1
4.4.1.1
4.5
4.5.1
4.5.1.1
4.5.1.2
4.5.1.3
4.5.1.4
4.5.1.5
4.5.2
4.5.3
4.5.4
4.5.5
4.5.6
4.5.7
4.5.7.1
4.5.7.2
4.5.8
4.5.8.1
4.5.8.2
4.5.8.3
4.5.8.4
4.5.8.5
4.5.8.6
4.5.9
4.5.9.1
4.5.9.2
4.5.9.3
4.5.9.4
4.5.9.5
4.5.9.5.1
4.5.9.5.2
4.5.9.5.3
4.5.9.5.4
4.5.9.5.5
4.5.9.5.6
4.5.9.5.7
4.6
4.6.1
4.6.1.1
Title
Page
Pitch axis control displacements -- takeoff. ........................................................................... 363
Pitch axis control displacements--maneuvering. ................................................................... 365
Flying quality requirements for normal (flight path) axis........................................................ 366
Flight path response to attitude change ................................................................................ 367
Transient flight path response to attitude change. ................................................................ 367
Steady-state flight path response to attitude change. ........................................................... 371
Flight path response to designated flight path controller....................................................... 381
Flight path control power ...................................................................................................... 382
Control power for designated primary flight path controller................................................... 382
Control power for designated secondary flight path controller. ............................................. 382
Flight path controller characteristics...................................................................................... 382
Flying qualities requirements for longitudinal (speed) axis ................................................... 384
Speed response to attitude changes..................................................................................... 384
Speed response to attitude changes--relaxation in transonic flight....................................... 389
Flying qualities requirements for the roll axis ........................................................................ 391
Roll response to roll controller ............................................................................................... 391
Roll mode............................................................................................................................... 391
Spiral stability......................................................................................................................... 406
Coupled roll-spiral oscillation................................................................................................. 410
Roll oscillations...................................................................................................................... 418
Roll time delay. ...................................................................................................................... 433
Pilot-induced roll oscillations. ................................................................................................ 436
Linearity of roll response o roll controller............................................................................... 437
Lateral acceleration at pilot station........................................................................................ 438
Roll characteristics in steady sideslip.................................................................................... 442
Roll axis control for takeoff and landing in crosswinds.......................................................... 444
Roll axis response to other inputs ......................................................................................... 445
Roll axis response to augmentation failures.......................................................................... 445
Roll axis response to configuration or control mode change. ............................................... 447
Roll axis control power .......................................................................................................... 448
Roll axis response to roll control inputs................................................................................. 448
Roll axis control power in steady sideslips. ........................................................................... 476
Roll axis control power in crosswinds.................................................................................... 477
Roll axis control power for asymmetric thrust........................................................................ 480
Roll axis control power in dives and pullouts......................................................................... 482
Roll axis control power for asymmetric loading. .................................................................... 483
Roll axis control forces and displacements ........................................................................... 484
Roll control displacements..................................................................................................... 488
Roll axis control forces to achieve required roll performance. .............................................. 490
Roll axis control sensitivity..................................................................................................... 492
Roll axis control centering and breakout forces. ................................................................... 510
Roll axis control force limits ................................................................................................... 511
Roll axis control force limits in steady turns. ......................................................................... 511
Roll axis control force limits in dives and pullouts. ................................................................ 513
Roll axis control force limits in crosswinds. ........................................................................... 514
Roll axis control force limits in steady sideslips..................................................................... 515
Roll axis control force limits for asymmetric thrust. ............................................................... 516
Roll axis control force limits for failures. ................................................................................ 517
Roll axis control force limits for configuration or control mode change................................. 518
Flying quality requirements for the yaw axis ......................................................................... 519
Yaw axis response to yaw and side-force controllers ........................................................... 519
Dynamic lateral-directional response. ................................................................................... 519
Supersedes pages iv - xiii of
MIL-STD-1797A
xiv
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
Title
Page
4.6.1.2
4.6.1.3
4.6.2
4.6.3
4.6.4
4.6.5
4.6.5.1
4.6.5.2.
4.6.5.3
4.6.6
4.6.6.1
4.6.6.2
4.6.6.3
4.6.7
4.6.7.1
4.6.7.2
4.6.7.3
4.6.7.4
4.6.7.5
4.6.7.6
4.6.7.7
4.6.7.8
4.6.7.9
4.6.7.10
4.6.7.11
4.7
4.7.1
4.8
4.8.1
4.8.2
4.8.3
4.8.4
4.8.4.1
4.8.4.2
4.8.4.2.1
4.8.4.2.2
4.8.4.2.3
4.8.4.2.4
4.8.4.3
4.8.4.3.1
4.8.4.3.2
4.9
4.9.1
4.9.2
4.9.3
Steady sideslips..................................................................................................................... 542
Wings-level turn..................................................................................................................... 545
Yaw axis response to roll controller....................................................................................... 554
Pilot-induced yaw oscillations................................................................................................ 584
Yaw axis control for takeoff and landing in crosswinds......................................................... 585
Yaw axis response to other inputs ........................................................................................ 588
Yaw axis response to asymmetric thrust............................................................................... 588
Yaw axis response to failures................................................................................................ 593
Yaw axis response to configuration or control mode change................................................ 594
Yaw axis control power.......................................................................................................... 595
Yaw axis control power for takeoff, landing, and taxi. ........................................................... 596
Yaw axis control power for asymmetric thrust....................................................................... 598
Yaw axis control power with asymmetric loading. ................................................................. 599
Yaw axis control forces.......................................................................................................... 600
Yaw axis control force limits in rolling maneuvers................................................................. 601
Yaw axis control force limits in steady turns.......................................................................... 602
Yaw axis control force limits during speed changes. ............................................................ 604
Yaw axis control force limits in crosswinds. .......................................................................... 605
Yaw axis control force limits with asymmetric loading. ......................................................... 606
Yaw axis control force limits in dives and pullouts. ............................................................... 607
Yaw axis control force limits for waveoff (go-around). .......................................................... 608
Yaw axis control force limits for asymmetric thrust during, takeoff. ...................................... 609
Yaw axis control force limits with flight control failures. ........................................................ 610
Yaw axis control force limits -- control mode change............................................................ 611
Yaw axis breakout forces. ..................................................................................................... 612
Flying qualities requirements for the lateral flight path axis .................................................. 613
Dynamic response for lateral translation. .............................................................................. 613
Flying qualities requirements for combined axes .................................................................. 616
Cross-axis coupling in roll maneuvers................................................................................... 616
Crosstalk between pitch and roll controllers.......................................................................... 620
Control harmony. ................................................................................................................... 622
Flight at high angle of attack.................................................................................................. 624
Warning cues......................................................................................................................... 630
Stalls. ..................................................................................................................................... 632
Stall approach........................................................................................................................ 634
Stall characteristics................................................................................................................ 637
Stall prevention and recovery. ............................................................................................... 639
One-engine-out stalls. ........................................................................................................... 643
Post-stall gyrations and spins................................................................................................ 645
Departure from controlled flight. ............................................................................................ 649
Recovery from post-stall gyrations and spins........................................................................ 657
Flying qualities requirements in atmospheric disturbances................................................... 663
Allowable flying qualities degradations in atmospheric disturbances. .................................. 663
Definition of atmospheric disturbance model form. ............................................................... 671
Application of disturbance models in analyses...................................................................... 684
5.1
5.1.1
5.1.2
5.1.3
5.1.4
5.1.4.1
General requirements-verification ......................................................................................... 102
Loadings-verification.............................................................................................................. 102
Moments and products of inertia-verification......................................................................... 103
Internal and external stores-verification. ............................................................................... 105
Flight Envelopes-verification.................................................................................................. 111
Operational Flight Envelopes-verification.............................................................................. 111
Supersedes pages iv - xiii of
MIL-STD-1797A
xv
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
5.1.4.2
5.1.4.3
5.1.5
5.1.6
5.1.6.1
5.1.6.2
5.1.6.3
5.1.7
5.1.7.1
5.1.7.2
5.1.7.3
5.1.7.4
5.1.7.5
5.1.7.6
5.1.8
5.1.8.1
5.1.8.2
5.1.9
5.1.10
5.1.11
5.1.11.1
5.1.11.2
5.1.11.3
5.1.11.4
5.1.11.5
5.1.11.6
5.1.11.7
5.1.11.8
5.1.12
5.1.12.1
5.1.12.2
5.1.12.3
5.1.12.4
5.1.12.5
5.1.12.6
5.1.12.7
5.1.12.8
5.1.12.9
5.1.12.10
5.1.12.11
5.1.13
5.1.13.1
5.1.13.2
5.1.13.3
5.1.3.4
5.1.13.5
5.1.13.6
5.2
5.2.1
5.2.1.1
5.2.1.2
5.2.2
Title
Page
Service Flight Envelopes-verification. ................................................................................... 114
Permissible Flight Envelopes-verification.............................................................................. 116
Configurations and States of the aircraft-verification. ........................................................... 118
Aircraft Normal States-verification......................................................................................... 119
Allowable Levels for Aircraft Normal States-verification........................................................ 121
Flight outside the Service Flight Envelopes-verification........................................................ 121
Ground operation-verification. ............................................................................................... 121
Aircraft Failure States-verification. ........................................................................................ 122
Allowable Levels for Aircraft Failure States-verification ........................................................ 125
Aircraft Special Failure States-verification............................................................................. 125
Probability calculation-verification. ........................................................................................ 131
Generic failure analysis-verification....................................................................................... 137
When Levels are not specified-verification............................................................................ 138
Failures outside the Service Flight Envelope-verification...................................................... 139
Dangerous flight conditions-verification................................................................................. 141
Warning and indication-verification. ...................................................................................... 141
Devices for indication, warning, prevention, and recovery-verification. ................................ 141
Interpretation of subjective requirements-verification............................................................ 142
Interpretation of quantitative requirements-verification. ........................................................ 145
General flying qualities-verification........................................................................................ 147
Buffet-verification................................................................................................................... 147
Release of stores-verification. ............................................................................................... 149
Effects of armament delivery and special equipment-verification. ........................................ 150
Failures-verification. .............................................................................................................. 154
Control margin-verification..................................................................................................... 158
Pilot-induced oscillations (PIO)-verification. .......................................................................... 163
Residual oscillations-verification. .......................................................................................... 165
Control cross-coupling-verification. ....................................................................................... 166
General flight control system characteristics-verification. ..................................................... 169
Control centering and breakout forces-verification................................................................ 169
Cockpit control free play-verification. .................................................................................... 169
Adjustable controls-verification.............................................................................................. 170
Rate of control displacement-verification. ............................................................................. 174
Dynamic characteristics-verification. ..................................................................................... 174
Damping-verification.............................................................................................................. 174
Transfer to alternate control modes-verification. ................................................................... 174
Flight control system failures-verification. ............................................................................. 174
Augmentation systems-verification........................................................................................ 174
Auxiliary dive recovery devices-verification........................................................................... 174
Direct force controllers-verification. ....................................................................................... 174
General trim requirements-verification .................................................................................. 178
Trim system irreversibility-verification. .................................................................................. 178
Rate of trim operation-verification. ........................................................................................ 178
Stalling of trim systems-verification. ...................................................................................... 178
Transients and trim changes-verification............................................................................... 178
Trim for asymmetric thrust-verification. ................................................................................. 178
Automatic trim system-verification......................................................................................... 178
Flying qualities requirements for the pitch axis-verification................................................... 181
Pitch attitude dynamic response to pitch controller-verification ............................................ 181
Long-term pitch response-verification. .................................................................................. 181
Short-term pitch response-verification................................................................................... 272
Pilot-induced pitch oscillations-verification. ........................................................................... 288
Supersedes pages iv - xiii of
MIL-STD-1797A
xvi
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
5.2.3
5.2.4
5.2.5
5.2.6.1
5.2.6.2
5.2.7.1
5.2.7.2
5.2.7.2.1
5.2.7.3
5.2.7.4
5.2.8.1
5.2.8.2
5.2.8.3
5.2.8.4
5.2.8.5
5.2.8.6.1
5.2.8.6.2
5.2.8.6.3
5.2.8.6.4
5.2.8.6.5
5.2.8.6.6
5.2.8.7
5.2.9
5.2.9.1
5.2.9.2
5.3
5.3.1
5.3.1.1
5.3.1.2
5.3.2
5.3.3
5.3.3.1
5.3.3.2
5.3.4
5.4
5.4.1
5.4.1.1
5.5
5.5.1
5.5.1.1
5.5.1.2
5.5.1.3
5.5.1.4
5.5.1.5
5.5.2
5.5.3
5.5.4
5.5.5
5.5.6
5.5.7
5.5.7.1
Title
Page
Residual pitch oscillations-verification................................................................................... 290
Normal acceleration at pilot station-verification..................................................................... 293
Pitch trim changes-verification............................................................................................... 295
Pitch axis response to failures, control free-verification. ....................................................... 297
Pitch axis response to configuration or control mode change -verification. .......................... 299
Pitch axis control power in unaccelerated flight-verification.................................................. 301
Pitch axis control power in maneuvering flight-verification.................................................... 303
Load factor response--verification. ........................................................................................ 305
Pitch axis control power in takeoff-verification. ..................................................................... 308
Pitch axis control power in landing-verification...................................................................... 310
Pitch axis control forces--steady-state control force per g-verification.................................. 329
Pitch axis control forces-transient control force per g-verification......................................... 339
Pitch axis control forces--control force variations during rapid speed changes
-verification. ........................................................................................................................... 340
Pitch axis control forces--control force vs. control deflection-verification.............................. 348
Pitch axis control breakout forces-verification. ...................................................................... 350
Pitch axis control force limits-takeoff-verification................................................................... 352
Pitch axis control force limits--landing -verification................................................................ 353
Pitch axis control force limits--dives-verification.................................................................... 355
Pitch axis control force limits--sideslips-verification. ............................................................. 357
Pitch axis control force limits--failures-verification................................................................. 359
Pitch axis control force limits--control mode change -verification. ........................................ 360
Pitch axis trim systems-verification. ...................................................................................... 362
Pitch axis control displacements-verification......................................................................... 363
Pitch axis control displacements--takeoff-verification. .......................................................... 363
Pitch axis control displacements--maneuvering-verification. ................................................ 365
Flying qualities requirements for the normal (flight path) axis-verification ............................ 369
Flight path response to attitude change-verification.............................................................. 369
Transient flight path response to attitude change-verification............................................... 369
Steady-state flight path response to attitude change-verification.......................................... 380
Flight path response to designated flight path controller-verification. ................................... 381
Flight path control power ....................................................................................................... 382
Control power for designated primary flight path controller-verification. ............................... 382
Control power for designated secondary flight path controller-verification. .......................... 382
Flight path controller characteristics-verification. .................................................................. 382
Flying qualities requirements for longitudinal (speed) axis-verification ................................. 386
Speed response to attitude changes-verification. ................................................................. 386
Speed response to attitude changes--relaxation in transonic flight-verification. ................... 389
Flying quality requirements for roll axis-verification .............................................................. 404
Roll response to roll controller-verification ............................................................................ 404
Roll mode-verification. ........................................................................................................... 404
Spiral stability-verification. ..................................................................................................... 409
Coupled roll-spiral oscillation-verification. ............................................................................. 416
Roll oscillations-verification. .................................................................................................. 431
Time delay-verification........................................................................................................... 435
Pilot-induced roll oscillations-verification............................................................................... 436
Linearity of roll response to roll controller-verification. .......................................................... 437
Lateral acceleration at the pilot station-verification. .............................................................. 440
Roll characteristics in steady sideslip-verification. ................................................................ 443
Roll axis control for takeoff and landing in crosswinds-verification. ...................................... 444
Roll axis response to other inputs-verification....................................................................... 445
Roll axis response to augmentation failures-verification. ...................................................... 445
Supersedes pages iv - xiii of
MIL-STD-1797A
xvii
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
5.5.7.2
5.5.8
5.5.8.1
5.5.8.2
5.5.8.3
5.5.8.4
5.5.8.5
5.5.8.6
5.5.9.1
5.5.9.2
5.5.9.3
5.5.9.4
5.5.9.5
5.5.9.5.1
5.5.9.5.2
5.5.9.5.3
5.5.9.5.4
5.5.9.5.5
5.5.9.5.6
5.5.9.5.7
5.6
5.6.1
5.6.1.1
5.6.1.2
5.6.1.3
5.6.2
5.6.3
5.6.4
5.6.5
5.6.5.1
5.6.5.2
5.6.5.3
5.6.6
5.6.6.1
5.6.6.2
5.6.6.3
5.6.7
5.6.7.1
5.6.7.2
5.6.7.3
5.6.7.4
5.6.7.5
5.6.7.6
5.6.7.7
5.6.7.8
5.6.7.9
5.6.7.10
5.6.7.11
5.7
5.7.1
5.8
5.8.1
Title
Page
Roll axis response to configuration or control mode change-verification.............................. 447
Roll axis control power-verification........................................................................................ 474
Roll axis response to roll control inputs-verification. ............................................................. 474
Roll axis control power in steady sideslips--verification. ....................................................... 476
Roll axis control power in crosswinds-verification. ................................................................ 478
Roll axis control power for asymmetric thrust -verification. ................................................... 480
Roll axis control power in dives and pullouts-verification. ..................................................... 482
Roll axis control power for asymmetric loading-verification. ................................................. 483
Wheel control displacements -verification. ............................................................................ 488
Roll axis control forces o achieve required roll performance-verification.............................. 491
Roll axis control sensitivity-verification. ................................................................................. 507
Roll axis control centering and breakout forces-verification.................................................. 510
Roll axis control force limits-verification ................................................................................ 511
Roll axis control force limits in steady turns-verification........................................................ 511
Roll axis control force limits in dives and pullouts-verification............................................... 513
Roll axis control force limits in crosswinds-verification.......................................................... 514
Roll axis control force limits in steady sideslips-verification. ................................................. 515
Roll axis control force limits for asymmetric thrust-verification.............................................. 516
Roll axis control force limits for failures-verification............................................................... 517
Roll axis control force limits for configuration or control mode change-verification. ............. 518
Flying qualities requirements for the yaw axis-verification .................................................... 539
Yaw axis response to yaw and side-force controllers-verification......................................... 539
Dynamic lateral-directional response-verification.................................................................. 539
Steady sideslips-verification. ................................................................................................. 543
Wings-level turn -verification. ................................................................................................ 553
Yaw axis response to roll controller-verification. ................................................................... 580
Pilot-induced yaw oscillations-verification. ............................................................................ 584
Yaw axis control for takeoff and landing in crosswinds-verification. ..................................... 586
Yaw axis response to other inputs-verification ...................................................................... 589
Yaw axis response to asymmetric thrust-verification. ........................................................... 589
Yaw axis response to failures-verification. ............................................................................ 593
Yaw axis response to configuration or control mode change-verification. ............................ 594
Yaw axis control power-verification. ...................................................................................... 595
Yaw axis control power for takeoff, landing, and taxi-verification.......................................... 597
Yaw axis control power for asymmetric thrust-verification. ................................................... 598
Yaw axis control power with asymmetric loading-verification. .............................................. 599
Yaw axis control forces-verification. ...................................................................................... 600
Yaw control force limits in rolling maneuvers-verification...................................................... 601
Yaw axis control force limits in steady turns-verification. ...................................................... 602
Yaw axis control force limits during speed changes-verification. .......................................... 604
Yaw axis control force limits in crosswinds-verification......................................................... 605
Yaw axis control force limits with asymmetric loading-verification. ....................................... 606
Yaw axis control force limits in dives and pullouts-verification.............................................. 607
Yaw axis control force limits for waveoff (go-around)-verification. ........................................ 608
Yaw axis control force limits for asymmetric thrust during takeoff-verification...................... 609
Yaw axis control force limits with flight control failures-verification. ...................................... 610
Yaw axis control force limits -- control mode change-verification. ........................................ 611
Yaw axis breakout forces-verification.................................................................................... 612
Flying qualities requirements for the lateral flight path axis-verification ................................ 615
Dynamic response for lateral translation-verification. ........................................................... 615
Flying qualities requirements for combined axes-verification................................................ 618
Cross-axis coupling in roll maneuvers-verification. ............................................................... 618
Supersedes pages iv - xiii of
MIL-STD-1797A
xviii
MIL-STD-1797A
APPENDIX A CONTENTS
Paragraph
Number
5.8.2
5.8.3
5.8.4
5.8.4.1
5.8.4.2
5.8.4.2.1
5.8.4.2.2
5.8.4.2.3
5.8.4.2.4
5.8.4.3
5.8.4.3.1
5.8.4.3.2
5.9
5.9.1
5.9.2
Title
Page
Crosstalk between pitch and roll controllers-verification. ...................................................... 620
Control harmony-verification.................................................................................................. 622
Flight at high angle of attack-verification. .............................................................................. 626
Warning cues-verification. ..................................................................................................... 631
Stalls-verification. .................................................................................................................. 633
Stall approach-verification. .................................................................................................... 635
Stall characteristics-verification. ............................................................................................ 638
Stall prevention and recovery-verification. ............................................................................ 640
One-engine-out stalls-verification.......................................................................................... 643
Post-stall gyrations and spins-verification. ............................................................................ 646
Departure from controlled flight-verification........................................................................... 650
Recovery from post-stall gyrations and spins-verification. .................................................... 661
Flying qualities requirements in atmospheric disturbances-verification ................................ 670
Allowable flying qualities degradations in atmospheric disturbances-verification. ................ 670
Definition of atmospheric disturbance model form-verification.............................................. 692
Supersedes pages iv - xiii of
MIL-STD-1797A
xix
MIL-STD-1797A
APPENDIX B CONTENTS
Paragraph
Title
Page
Number
Determining Equivalent Systems .......................................................................................... 688
APPENDIX C CONTENTS
Paragraph
Title
Page
Number
Background Information and Cross Reference ..................................................................... 692
Numerical cross-index of MIL-F-8785C to Appendix A ......................................................... 699
Supersedes pages iv - xiii of
MIL-STD-1797A
xx
MIL-STD-1797A
FIGURES
Figure
Standard
Page
FIGURE 1. Roll-sideslip coupling parameters--right rolls ..................................................................... 10/11
FIGURE 2. Roll-Sideslip coupling parameters--left rolls. - continued ................................................... 12/13
Supersedes pages iv - xiii of
MIL-STD-1797A
xxi
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 3.
FIGURE 4.
FIGURE 5.
FIGURE 6.
FIGURE 7.
FIGURE 8.
FIGURE 9.
FIGURE 10.
FIGURE 11.
FIGURE 12.
FIGURE 13.
FIGURE 14.
FIGURE 15.
FIGURE 16.
FIGURE 17.
FIGURE 18.
FIGURE 19.
FIGURE 20.
FIGURE 21.
FIGURE 22.
FIGURE 23
FIGURE 24.
FIGURE 25.
FIGURE 26.
FIGURE 27.
FIGURE 28.
FIGURE 29.
FIGURE 30.
FIGURE 31.
FIGURE 32.
FIGURE 33.
FIGURE 34.
FIGURE 35.
FIGURE 36.
FIGURE 37.
FIGURE 38.
Page
Classification of aircraft (AFFDL-TR-69-72). .................................................................... 84
Ground attack maneuver scenario.................................................................................... 87
Response to a 1-second, half-stick aileron doublet for tracking scenario. ....................... 88
Definition of Flying Quality Levels in Calm to Light Turbulence. ...................................... 91
Definition of Flight Envelope terms. .................................................................................. 99
Definition of Levels which include atmospheric disturbances as
well as failures suggested by Carlson (AFFDL-TR-78-171). .......................................... 127
Buffet intensity rise determination from NASA-TP-1368................................................. 141
Control surface requirements.......................................................................................... 153
Control margin requirements. ......................................................................................... 155
PIO tendency classification from AFWAL-TR-81-3118................................................... 157
Short-period dynamic requirements................................................................................ 179
Effect of fitting with 1/ Tθ2 fixed and free, Category C requirements............................... 183
Envelopes of maximum unnoticeable added dynamics
(AFWAL-TR-82-3064)..................................................................................................... 186
Comparison of Bode amplitude plots for basic and augmented
configurations of table XIII. ............................................................................................. 188
Comparison of phase angles for basic and augmented
configurations of table XIII. ............................................................................................. 188
Effect of first- and second-order lags on equivalent time delay
and pilot rating: LAHOS configurations (AFFDL-TR-78-122). ........................................ 189
Pilot control of pitch attitude and flight path. ................................................................... 190
Definition of CAP from frequency response asymptotes. ............................................... 192
Pitch acceleration response to a unit step force input. ................................................... 194
Time delay versus CAP’ - Neal-Smith data (from NADC-81186-60). ............................. 195
Time delay versus CAP’- LAHOS data (from NADC-81186-60)..................................... 195
Requirements for short-term pitch response to pitch controller
(ωspTθ2 vs ζsp). ................................................................................................................. 197
Comparison of pilot ratings with category A short-period frequency
requirements. .................................................................................................................. 200
Comparison of pilot ratings with Category A short-period damping
requirements. .................................................................................................................. 201
Comparison of pilot ratings with Category A short-period frequency
and damping ratio requirements. .................................................................................... 202
Category A C-5A flight test data. .................................................................................... 203
Alternate Category A flying qualities requirements for short-period
pitch response................................................................................................................. 203
Comparison of pilot ratings with Category B short-period frequency
requirements. .................................................................................................................. 205
Comparison of pilot ratings with Category B short-period damping
requirements. .................................................................................................................. 205
Category B short-period characteristics.......................................................................... 206
L-1011 climb short-period characteristics. ...................................................................... 206
L-1011 cruise short-period characteristics (AFWAL-TR-83-3015). ................................ 206
L-1011 descent short-period characteristics (AFWAL-TR-83-3015). ............................. 206
Alternative Category B short-period flying qualities requirements
(NASA-TM-X-1584 data, Level 1 Fs/n)............................................................................ 207
Comparison of pilot ratings with Category C short-period frequency
requirements. .................................................................................................................. 208
Comparison of pilot ratings with Category C short-period damping
requirements. .................................................................................................................. 209
Supersedes pages iv - xiii of
MIL-STD-1797A
xxii
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 39.
FIGURE 40.
FIGURE 41.
FIGURE 42.
FIGURE 43.
FIGURE 44.
FIGURE 45.
FIGURE 46.
FIGURE 47.
FIGURE 48.
FIGURE 49.
FIGURE 50.
FIGURE 51.
FIGURE 52.
FIGURE 53.
FIGURE 54.
FIGURE 55.
FIGURE 56.
FIGURE 57.
FIGURE 58.
FIGURE 59.
FIGURE 60.
FIGURE 61a.
FIGURE 61b.
FIGURE 62.
FIGURE 63.
FIGURE 64.
FIGURE 65.
FIGURE 66.
FIGURE 67.
FIGURE 68.
FIGURE 69.
Page
Pilot ratings for large airplanes (nominal equivalent short-period
parameters) from AFWAL-TR-81-3118........................................................................... 210
Category C flight data for the Lockheed C-5A and L-1011,
AFWAL-TR-83-3015. ...................................................................................................... 211
Tlme delay bands associated with flying qualities boundaries vs
bandwidth, AFWAL-TR-81-3118..................................................................................... 212
Alternate Category C short-period flying qualities requirements. ................................... 213
Comparison of LOES dynamics with short-period requirements;
Category A, Neal-Smith (AFFDL-TR-70-74) configurations, MCAIR
(MDC Rpt A6792) matches............................................................................................. 215
Comparison of Neal-Smith LOES characteristics with ωspTθ2 vs ζsp ............................... 216
Comparison of LOES dynamics with short-period requirements;
Category C, LAHOS (AFFDL-TR-78-122) configurations, MCAIR
("Equivalent Systems Approach for Flying Qualities Specification"). ............................ 218
Comparison of LAHOS LOES characteristics with ωsp Tθ2 vs. ζsp................................... 219
Comparison of effects of various stability characteristics on
airplane response to elevator pulse (-5 deg for 0.2 sec at t = 0)
(NASA-TM-X-62)............................................................................................................. 220
Contours of constant pilot opinion in statically unstable region;
constant stick-to-stabilizer gearing (NASA-TN-D-779). .................................................. 221
Pitch rate response to step input of pitch controller force or deflection. ......................... 224
Comparison of YF-12 and XB-70 handing qualities evaluation with
the GPAS results (NASA-CR-159236)............................................................................ 227
Short-period frequency (NASA-CR-159236). ................................................................. 227
Pilot rating correlation with effective time delay (AFFDL-TR-78-122 data). ................... 228
Pilot rating with effective time delay (AFWAL-TR-81-3116 data). .................................. 228
Pilot rating correlation with effective time delay (AFFDL-TR-78-122). ........................... 228
Pilot rating correlation with effective time delay (AFFDL-TR-68-90 data). ..................... 228
230
Pilot rating correlation with effective time delay (AFFDL-TR-70-74 data). ..................... 230
Bandwidth requirements. ................................................................................................ 231
Simplified pilot-vehicle closure for pitch control. ............................................................. 233
Definition of bandwidth frequency ωBW from open loop frequency response. ................ 235
Comparison of Neal-Smith data (AFFDL-TR-70-74) with bandwidth
(mean ratings). ................................................................................................................ 235
Level 1/2 system of Neal-Smith (lD): ωBW = 2.7 rad/sec, mean PR = 4.1...................... 236
Level 3 system of Neal-Smith (2I) ωBW = 2.5 mean PR = 8.0. ....................................... 236
Correlation of pilot ratings with ωBW and τe (AFFDL-TR-70-74 data).............................. 237
Large difference in bandwidth due to shelf in amplitude plot with
moderate values of τp (configurations of AFFDL-TR-78-122)......................................... 238
Correlation of pilot ratings with ωBW and τp for Neal-Smith data
(Category A) (data from AFFDL-TR-70-74, ratings in parentheses
from AFFDL-TR-74-9)..................................................................................................... 239
Correlation of pilot ratings with ωBW and τp for approach and
landing (AFFDL-TR-122). ............................................................................................... 240
Comparison of pilot ratings for Category A short-period
configurations with bandwidth (classical airplanes). ....................................................... 241
Comparison of pilot ratings for Category C short-period
configurations with bandwidth (classical airplanes). ....................................................... 242
Design criteria for pitch dynamics with the pilot in the loop. ........................................... 244
Amplitude-phase plot for configuration 13 (fwd; c.g.). .................................................... 247
Supersedes pages iv - xiii of
MIL-STD-1797A
xxiii
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 70.
FIGURE 71.
FIGURE 72.
FIGURE 73.
FIGURE 74.
FIGURE 75.
FIGURE 76.
FIGURE 77.
FIGURE 78.
FIGURE 79.
FIGURE 80.
FIGURE 81.
FIGURE 82.
FIGURE 83.
FIGURE 84.
FIGURE 85.
FIGURE 86.
FIGURE 87.
FIGURE 88.
FIGURE 89a.
FIGURE 89b.
FIGURE 90.
FIGURE 91.
FIGURE 92.
FIGURE 93.
FIGURE 94.
FIGURE 95.
FIGURE 96.
FIGURE 97.
FIGURE 98.
FIGURE 99.
FIGURE 100.
FIGURE 101.
FIGURE 102.
FIGURE 103.
FIGURE 104.
FIGURE 105.
FIGURE 106.
FIGURE 107.
Page
Amplitude-phase plot for configuration 12 (aft; c.g. - unstable). ..................................... 247
Amplitude-phase plot for configuration 6-1. .................................................................... 248
Amplitude-phase plot for configuration 6-2. .................................................................... 248
Pitch short period time responses................................................................................... 251
Equivalent ωn boundaries................................................................................................ 252
Design aim criteria for pitch attitude frequency response............................................... 252
High and low order frequency response. ........................................................................ 253
Trends of high order phase rate...................................................................................... 253
Flight path - attitude relationships. .................................................................................. 254
nz hang-on effects. .......................................................................................................... 254
Precision tracking: q ∼ θ trends....................................................................................... 255
New short-period thumbprint (from ICAS-86-5.3.4). ....................................................... 255
Low-order pilot - aircraft attitude frequency response. ................................................... 256
Comparison of equivalent delay effects in pitch or roll response
to stick force for different simulations (from AIAA Paper 80-1611-CP). ......................... 260
AFTI/F-16 independent back-up pitch rate feedback block diagram. ............................. 261
AFTI/F-16 q → δe feedback (IBU). .................................................................................. 262
AFTI/F-16 θ → Fs for IBU (q → δe closed). ..................................................................... 263
Frequency and time response comparison..................................................................... 264
Maximum pitch rate overshoot for a step control input (from
DOT/FAA/CT-82/130-II). ................................................................................................. 266
Typical DFC control frequency sweep. ........................................................................... 269
Fourier transformed heading response........................................................................... 269
Approach and landing, no pilot-induced oscillation, configuration
P12 of AFWAL-TR-81-3116, medium offset approach (75 ft lateral,
50 ft vertical), landing no. 1 (from AIAA Paper 80-1611-CP).......................................... 270
Pilot-induced oscillation at touchdown. ........................................................................... 271
YF-17 pitch attitude dynamics (AFFDL-TR-77-57). ........................................................ 277
YF-17 acceleration control system dynamics (AFFDL-TR-77-57).................................. 278
Effect of dither on B-1 limit cycle oscillations (from AFFTC-TR-79-2). ........................... 283
Pilot rating vs pilot position - center of rotation (from AFWAL-TR-81-3118). ................. 285
Nosewheel and tailwheel lift-off. ..................................................................................... 299
Effect of arm/stick geometry on maximum push and pull capability
by the right arm for the 5th percentile male (Human Engineering
Guide to Equipment Design)........................................................................................... 304
Effect of upper arm angle on pull and push strength for the 5th and
95th percentile male (Human Engineering Guide to Equipment Design). ...................... 305
Effect of arm position and wheel angle on maximum push and pull
capability by the right arm for the 5th percentile male (Human
Engineering Guide to Equipment Design). ..................................................................... 306
Elevator maneuvering force gradient limits: center-stick controller, nL 7.0.................... 311
Elevator maneuvering force gradient limits: wheel controller, nL = 3.0.......................... 311
OV-10A maneuvering control (AFFDL-TR-78-171). ....................................................... 312
Longitudinal stick force at stall (AFFDL-TR-78-171)....................................................... 312
Short-period frequency vs longitudinal stick force per g (FS /δ
separately optimized) for a side-stick controller.............................................................. 313
Comparison of optimum FS/n with limits of table XVII (AFFDL-TR66-163, Category A; nL = 7g). ......................................................................................... 314
Comparison of optimum FS/n with limits of table XVII (FDL-TDR64-60, Category C; nL = 3g). ........................................................................................... 315
Comparison of optimum FS/n from data of Neal and Smith
(AFFDL-TR-70-74) with limits of table XVII (Category A; nL = 7g). ................................ 316
Supersedes pages iv - xiii of
MIL-STD-1797A
xxiv
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 108.
FIGURE 109.
FIGURE 110.
FIGURE 111.
FIGURE 112.
FIGURE 113.
FIGURE 114.
FIGURE 115.
FIGURE 116.
FIGURE 117.
FIGURE 118.
FIGURE 119.
FIGURE 120.
FIGURE 121.
FIGURE 122.
FIGURE 123.
FIGURE 124.
FIGURE 125.
FIGURE 126.
FIGURE 127.
FIGURE 128.
FIGURE 129.
FIGURE 130.
FIGURE 131.
FIGURE 132.
FIGURE 133.
FIGURE 134.
FIGURE 135.
FIGURE 136.
FIGURE 137.
FIGURE 138.
FIGURE 139.
Page
Comparison of optimum FS/n from LAHOS (AFFDL-TR-78-122)
data with limits of table XIV (Category C; nL = 7g).......................................................... 317
Average pilot ratings for 1-lb/g segments of FS/n. .......................................................... 317
Comparison of FS/n with limits of table XVII, wheel controllers
(AFFDL-TR-68-91, Category A; nL = 3g). ....................................................................... 318
Elevator control force gradients for transport aircraft (from AFFDL-TR-78-171)............ 319
Comparison of FS/n for various Class I airplanes in landing approach
(Category C) with limits of table XVII. ............................................................................. 321
Illustration of resonance dip in FS/n due to low ζsp.......................................................... 325
The ratio (FS/n)/(FS/n)min vs. ζsp....................................................................................... 326
Sketch of effect of control system on resonant dip. ........................................................ 326
Comparison of (FS/n)min boundaries with (FS/n)SS and ζsp for cases
where (ωsp << ωcs (AFFDL-TR-66-163)........................................................................... 327
PIO tendency rating scale............................................................................................... 328
T-33 data from AFFDL-TR-70-74 (Level 1 equivalent ωsp, τe). ....................................... 329
PIO characteristics of A4D-2, T-38A, and F-4C (Douglas Aircraft Co.
LB-25452, FTC-TR-67-19, and “Investigation and Elimination of PIO
Tendencies in the Northrop T-38A”). .............................................................................. 330
PIO characteristics of airplanes described in AGARD-CP-17. ....................................... 330
Pilot comments for air-to-air tasks with standard harmony
(from AFFDL-TR-79-3126).............................................................................................. 334
Average pilot ratings for gross acquisition task. ............................................................. 337
Average pilot ratings for fine tracking tasks. ................................................................... 337
Control force per control displacement, Category A Flight Phases
--centerstick (from AFFDL-TR-69-72). ............................................................................ 339
Landing approach (T-33, AFFDL-TR-66-2). ................................................................... 367
Landing approach (T-33, AFFDL-TR-66-2). ................................................................... 367
Landing approach (T-33, AFFDL-TR-66-2). ................................................................... 368
Landing approach (T-33, AFFDL-TR-66-2). ................................................................... 368
Carrier approach (Ground simulator experiment, "Simulator and
Analytical Studies of Fundamental Longitudinal Control Problems
in Carrier Approach"). ..................................................................................................... 370
SST landing approach (Ground simulator experiments,
NASA-TN-D-2251 and AFFDL-TR-65-227). ................................................................... 370
Landing approach (AVRO 707, AGARD Rpt 420). ......................................................... 371
Ratings versus roll damping - flight test, moving-base, fixed base
with random input (from AFFDL-TR-65-138). ................................................................ 384
Proposed roll performance requirements (MIL-F-8785) for
Class III aircraft (from NADC-ED-6282).......................................................................... 386
Lateral control boundaries (from Princeton Univ Rpt 727). ............................................ 386
Lateral flying qualities boundaries (Lβ vs. TR , ζd = 0.1)
(from Princeton Univ Rpt 727). ....................................................................................... 387
Lateral flying qualities boundaries (Lβ vs. TR, ζd = 0.4)
(from Princeton Univ Rpt 727). ....................................................................................... 387
Pilot ratings and optimum aileron sensitivity (Medium φ/βd, Long TR )
(from AFFDL-TR-67-98).................................................................................................. 388
Variation of pilot opinion with Lδaδamax, for constant values of TR
as obtained from the stationary flight simulator (from NASA Memo 1-29-59A).............. 389
Variation of pilot opinion with Lδaδamax, for constant values of TR
as obtained from the moving flight simulator (from NASA Memo 1-29-59A).................. 389
Supersedes pages iv - xiii of
MIL-STD-1797A
xxv
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 140.
FIGURE 141.
FIGURE 142.
FIGURE 143.
FIGURE 144.
FIGURE 145.
FIGURE 146.
FIGURE 147.
FIGURE 148.
FIGURE 149.
FIGURE 150.
FIGURE 151.
FIGURE 152.
FIGURE 153.
FIGURE 154.
FIGURE 155.
FIGURE 156.
FIGURE 157.
FIGURE 158.
FIGURE 159.
FIGURE 160.
FIGURE 161.
FIGURE 162.
FIGURE 163.
FIGURE 164.
FIGURE 165.
FIGURE 166.
FIGURE 167.
FIGURE 168.
FIGURE 169.
FIGURE 170.
FIGURE 171.
FIGURE 172.
FIGURE 173.
Page
Comparison of pilot opinion boundaries obtained from the fixed and
moving flight simulators (from NASA Memo 1-29-59A) .................................................. 390
Range of parameters Lδ a δa max and TR covered in the flight investigation,
shown in comparison with the motion simulator drived boundaries
(from NASA Memo 1-29-59A)......................................................................................... 390
Comparison of in-flight pilot-opinion rating with those predicted
from flight simulator boundaries (from NASA Memo 1-29-59A). .................................... 391
Pilot rating versus roll mode time constant (from WADD-TR-61-147)............................ 392
Average pilot rating of roll mode time constant (from AFFDL-TR-65-39). ...................... 392
Effect of roll mode LATHOS (AFWAL-TR-81-3171), Category A. .................................. 394
Comparison of models and data for closed-loop stick deflection
responses under lateral vibration (AMRL-TR-73-78)...................................................... 394
Effect of 1/TR on high-frequency gain. ............................................................................ 395
Limits of satisfactory and tolerable rates of spiral divergence
(from Cornell Aero Lab TB-574-F-6). .............................................................................. 399
Data for all types of flying pilot opinion versus spiral damping
(from Cornell Aero Lab TB-1094-F-1). ............................................................................ 400
Composite pilot ratings for spiral descent of simulated reentry
vehicle (from NASA-CR-778).......................................................................................... 404
Composite pilot ratings for up-and-away flight; moderate |φ/β|d
(from NASA-CR-778). ..................................................................................................... 405
Composite pilot ratings for up-and-away flight; large |φ/β|d
(from NASA-CR-778). ..................................................................................................... 405
Pilot ratings for ground simulation of NASA-TN-D-5466 (Dutch
roll characteristics vary). ................................................................................................. 406
Pilot ratings for ground simulation of AFFDL-TR-65-39 [(ωφ/ωd)2=
0.64 - 1.10]. ..................................................................................................................... 406
Coupled roll-spiral mode characteristics for M2-F2 and M2-F3
lifting bodies (from NASA-TN-D-6496)............................................................................ 407
Roll rate oscillation limitations......................................................................................... 411
Bank angle oscillation limitations. ................................................................................... 411
Effect of relative pole/zero location on piloted control of bank angle.............................. 414
Locations of ωφ, zero corresponding to Category A and C and
Level 1 and 2 boundaries on figure 156 (TR = 0.5 sec, TS = ∞)
(from AFFDL-TR-69-72).................................................................................................. 415
Posc/Pav, as a function of the ratio of dutch roll period and spiral
root time constant (from AFFDL-TR-72-41).................................................................... 416
Flight phase Category A data, moderate |φ/β|d (from AFFDL-TR-67-98). ...................... 418
Flight phase Category A data, large and small |φ/β|d (from AFFDL-TR-67-98). ............. 418
Flight phase Category B data (from NASA-CR-778). ..................................................... 419
Flight phase Category B data (from WADD-TR-61-147). ............................................... 419
Flight phase Category C data (from Princeton Univ Rpt 727). ....................................... 420
Category C data (approach and wave-off); Cooper-Harper pilot
ratings (from AFFDL-TR-70-145).................................................................................... 420
Positive and negative dihedral data of Princeton Univ Rpt 604...................................... 421
Effect of time delay, LATHOS data (AFWAL-TR-81-3171). ........................................... 426
Pilot rating vs. time delay -- lateral-directional. ............................................................... 426
Lateral acceleration criterion versus pilot rating from NASA-CR-159236....................... 431
Pilot rating vs lateral acceleration criteria. ...................................................................... 432
Roll control effectiveness parameters for Class III aircraft, Category C ......................... 447
C-5A flight test data (from AFFDL-TR-75-3)................................................................... 448
Supersedes pages iv - xiii of
MIL-STD-1797A
xxvi
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 174.
FIGURE 175.
FIGURE 176.
FIGURE 177.
FIGURE 178.
FIGURE 179.
FIGURE 180.
FIGURE 181.
FIGURE 182.
FIGURE 183.
FIGURE 184.
FIGURE 185.
FIGURE 186.
FIGURE 187.
FIGURE 188.
FIGURE 189.
FIGURE 190.
FIGURE 191.
FIGURE 192.
FIGURE 193.
FIGURE 194.
FIGURE 195.
FIGURE 196.
FIGURE 197.
FIGURE 198.
FIGURE 199.
FIGURE 200.
FIGURE 201.
FIGURE 202.
FIGURE 203.
Page
Roll performance for Class III aircraft (from AFFDL-TR-78-171).................................... 450
Comparison of pilot ratings for Class III aircraft in Category B
Flight Phases with requirements of table XXIX (NASA-TN-D-5957). ............................. 452
Time to bank, 300 for CV-990 (NASA-TN-D-6811). ....................................................... 453
F-4 roll control effectiveness, time-to- bank 90°, CR configuration
(from AFFFDL-TR-70-155). ............................................................................................ 454
F-4 roll control effectiveness, time-to- bank 90°, CR configuration
(from AFFFDL-TR-70-155). ............................................................................................ 454
F-4 roll control effectiveness; CO configuration. Limits shown for
speed range M, table XXXII (from AFFDL-TR-70-155). ................................................. 455
F-5E roll performance at 0.8 nL, configuration CO (from
AFFDL-TR-78-171). ........................................................................................................ 456
F-14A rolling performance in configuration PA; DLC on (from Navy
Rpt No. SA-C7R-75). ...................................................................................................... 457
F-15C aileron roll characteristics (from AFFTC-TR-80-23). ........................................... 457
Time to roll 90° versus match for F/A-18A (Navy Rpt No. SA-14R-81) .......................... 458
Time to roll 360° versus match for F/A-18A (Navy Rpt No. SA-14R-81) ........................ 458
Roll performance characteristics in configuration PA (from Navy Rpt
No. SA-14R-81)............................................................................................................... 461
YF-16 rolling performance in cruise configuration; AFFTC-TR-75-15. ........................... 462
Roll performance summary (from AFFTC-TR-79-10). .................................................... 463
Roll performance summary (from AFFTC-TR-79-10). .................................................... 465
Effect of arm/stick geometry on maximum applied force to the
left and to the right by the right arm for the 5th percentile male
(Human Engineering Guide to Equipment Design)......................................................... 477
Effect of upper arm angle on maximum applied force to the left
and to the right for the 5th and 95th percentile male (from Human
Engineering Guide to Equipment Design) ...................................................................... 478
Effect of arm position and wheel angle on maximum applied force to
to the left and to the right for the 5th percentile male (Human
Engineering Guide to Equipment Design). ..................................................................... 479
Variation of pilot rating with bank angle in the first second for four
values of effective angle (from (NASA-CR-635) ............................................................. 481
Block diagram representation of full-authority roll rate command
augmentation systems. ................................................................................................... 485
Range of acceptable nonlinear roll command shaping networks based
on flight tests (Class IV aircraft, Flight Phase Category A, right roll).............................. 486
Comparison of Pmax/Fas for several conventional Class IV aircraft
with CAS curves of figure 194......................................................................................... 487
Evolution of the F-16 CAS shaping network. .................................................................. 488
YF-16 PIO due to excessive lateral stick sensitivity (from
AFFTC-TR-75-15). .......................................................................................................... 489
Roll ratchet during banking maneuvers (DIGITAC,
AFFTC-TR-76-15) h = 20,000 ft. M = 0.75. .................................................................... 491
Evolution of roll CAS network for YA-7D DIGITAC
(AFFTC-TR-76-15).......................................................................................................... 492
Steady rolls on YF-16 (AFFTC-TR-75-15)...................................................................... 493
Roll ratcheting experienced on LATHOS (AFWAL-TR-81-3171)
Configuration 5-2............................................................................................................. 495
Roll gradients for LATHOS configurations 5-2 and 5-3 (TR = 0.15
sec) compared with acceptable range from figure 194................................................... 496
Influence of prefilter lag on pilot ratings (AFWAL-TR-81-3171).
Supersedes pages iv - xiii of
MIL-STD-1797A
xxvii
MIL-STD-1797A
APPENDIX A FIGURES
Figure
Page
FIGURE 212.
TR = 0.15 sec................................................................................................................... 496
Typical roll sensitivity - high gain. ................................................................................... 497
Flight Phase Category A - force sensitivity. .................................................................... 498
Pilot ratings from NASA Memo 1-29-59A (Category A Flight Phase)............................ 500
Flight Phase Category C - force sensitivity (from Princeton Univ
Rpt 727). ......................................................................................................................... 501
Effect Of ζd on pilot ratings for in-flight and fixed-base simulations
of NASA-TN-D-1141ωd = 1.78 - 1.90 rad/sec (Category B)............................................ 514
Dutch roll data (from Cornell Aero Lab Rpt TB-574-F-3)................................................ 514
Dutch roll data (from IAS Paper 60-18). ......................................................................... 515
Dutch roll data from fixed-base simulation of re-entry task
(ASD-TDR-61-362) TR ≈ 0.40 sec, TS ≈ ∞. ..................................................................... 516
Dutch roll data (from NASA-CR-778; low ω 2d φ / β d ....................................................... 518
FIGURE 213.
Dutch roll data (from NASA-CR-778; high ω 2d φ / β d ...................................................... 519
FIGURE 214.
FIGURE 215.
Dutch roll data (from WADD-TR-61-147)........................................................................ 520
Data for high ω 2d φ / β d (from WADD-TR-61-147). ......................................................... 520
FIGURE 216.
FIGURE 217.
Dutch roll data from in-flight simulation of AFFDL-TR-67-98.......................................... 521
Dutch roll data from flight tests of NASA Memo 12-10-58A (F-86E;
low-frequency data). ....................................................................................................... 522
Dutch roll data from flight tests of NASA Memo 12-10-58A (F-86E;
high-frequency data). ...................................................................................................... 522
Dutch roll data from in-flight simulation of Princeton Univ Rpt 797
(Navion; pilot ratings shown for optimum values of Lβ)................................................... 523
Dutch roll data from In-flight simulation of Class II-L airplanes in
landing approach (T-33; AFFDL-TR-70-145).................................................................. 524
Dutch roll data from moving-base - bank simulator tests of NASATN-D-3910 (supersonic transport). ................................................................................. 524
Variation of pilot rating with ωφ/ωd for moving-base simulator and
flight data ζd ≈ 0.15 (NASA-TN-D-3910). ........................................................................ 525
∆ζω/ ∆( ω 2d |φ/β|d) required to maintain a given basic rating (from
AFFDL-TR-70-145), based upon data of figures 224 and 225. ...................................... 526
Dutch roll data from controls-fixed rudder kicks of WADC-TR-52-298
compared to limits of table XL......................................................................................... 527
Data upon which relation of figure 223 are based (controls-fixed
rudder kicks, WADC-TR-52-298; figure reproduced from AFFDL-TR-65-138). ............. 528
Dutch roll data on existing airplanes (from AGARD-CP-17)........................................... 529
Lateral directional damping for some Class III airplanes (from
AFFDL-TR-78-171). ........................................................................................................ 530
Finding the dutch roll envelope. ...................................................................................... 533
Definition of bandwidth frequency................................................................................... 540
Effect of DFC manipulator sensitivity configuration WLT1 (very low
coupling) (from AFWAL-TR-81-3027). ............................................................................ 540
Effect of DFC manipulator sensitivity configuration WLT5 (high
favorable yaw coupling) (from AFWAL-TR-81-3027)...................................................... 541
Effect of DFC manipulator sensitivity -- configuration WLT12 (very
high favorable roll coupling) (from AFWAL-TR-81-3027). .............................................. 542
Correlation of pilot ratings with heading bandwidth; wings-level turn
mode; air-to-air tracking task. ......................................................................................... 543
Correlation of pilot ratings with heading bandwidth for conventional
aircraft; ILS approach task. ............................................................................................. 544
FIGURE 204.
FIGURE 205.
FIGURE 206.
FIGURE 207.
FIGURE 208.
FIGURE 209.
FIGURE 210.
FIGURE 211.
FIGURE 218.
FIGURE 219.
FIGURE 220.
FIGURE 221.
FIGURE 222.
FIGURE 223.
FIGURE 224.
FIGURE 225.
FIGURE 226.
FIGURE 227.
FIGURE 228.
FIGURE 229.
FIGURE 230.
FIGURE 231.
FIGURE 232.
FIGURE 233.
FIGURE 234.
Supersedes pages iv - xiii of
MIL-STD-1797A
xxviii
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 235.
FIGURE 236.
FIGURE 237.
FIGURE 238.
FIGURE 239.
FIGURE 240.
FIGURE 241.
FIGURE 242.
FIGURE 243.
FIGURE 244.
FIGURE 245.
FIGURE 246.
FIGURE 247.
FIGURE 248.
Page
Sideslip excursion limitations for small roll control commands....................................... 546
Pilot ratings and optimum aileron sensitivity (low |φ/β|d, medium TR)
(AFFDL-TR-67-98). ......................................................................................................... 548
Flight Phase Category A data from AFFDL-TR-67-98.................................................... 549
Flight Phase Category A data from AFFDL-TR-67-98
( ∆β / k, ψβ from AFFDL-TR-72-41). ................................................................................ 551
∆βmax/k versus ψβ for evaluation points that meet Level 1posc/pav criteria,
Category A data (from AFFDL-TR-72-36). ..................................................................... 552
Category B data of WADD-TR-61-147............................................................................ 553
Category C configurations of Princeton Univ Rpt 727 (∆β/k, ψβ from
AFFDL-TR-72-41). .......................................................................................................... 554
∆βmax/k versus ψβ for configurations that meet Level 1 posc/pav, ζd, and
ζdωd criteria (from AFFDL-TR-70-145)............................................................................ 554
C-5A flight test data (from AFFDL-TR-75-3)................................................................... 556
Sideslip excursion data for Class III aircraft in Category B Flight Phases
(from AFFDL-TR-78-171)................................................................................................ 557
Sideslip excursions for F4H-1 airplane (from AFFDL-TR-70-155). ................................ 558
F-5A flight test data. ........................................................................................................ 559
Crossfeed parameter boundaries. .................................................................................. 560
Pilot rating boundaries for acceptable roll control in turbulence with
r&o p& o ≤ 0.03 (from Princeton Univ Rpt 797)................................................................. 562
FIGURE 249.
FIGURE 250.
FIGURE 251.
Bode asymptotes and time response of crossfeed........................................................ 565
Effect of removing high-frequency roots from numerators. ............................................ 568
Required crossfeed for r&o = 0......................................................................................... 569
FIGURE 252.
& o is small................................................................. 570
Pilot rating correlations when r&o p
FIGURE 253.
FIGURE 254.
Pilot rating correlation with crossfeed parameters.......................................................... 573
Required aileron-rudder sequencing for several operational aircraft,
SAS/CAS ON (from Systems Technology, Inc., TR-1090-1).......................................... 574
Time history of aft-c.g. deep stall encountered by F-16B ............................................... 633
(AFFTC-TR-79-18).......................................................................................................... 633
Anti-spin SAS for F-16B (α ≥ 29 deg). ............................................................................ 634
Modified departure rating scale (MOD II)........................................................................ 639
Dynamic stability design guide suggested by Titiriga ..................................................... 644
(AGARD-CP-199)............................................................................................................ 644
Departure and spin susceptibility criterion suggested by ASD-TR-72-48
and AFWAL-TR-80-3141. ............................................................................................... 645
Departure susceptibility rating versus lateral closed-loop divergence
potential (from AFWAL-TR-80-3141). ............................................................................. 647
Left flat spin, F-4B (from AFFDL-TR-70-155). ................................................................ 652
Turbulence severity and exceedance probability............................................................ 658
Probability of exceeding mean wind speed at 20 ft. ....................................................... 660
Earth-axis winds.............................................................................................................. 666
Magnitude of discrete gusts. ........................................................................................... 668
Low-altitude turbulence integral scales........................................................................... 670
Horizontal turbulence RMS intensities. ........................................................................... 670
CVA ship burble steady wind ratios. ............................................................................... 672
u-component burble time constant and variance............................................................ 674
Comparative approximate frequency regimes of mission/aircraft-centered
and atmospheric disturbance features............................................................................ 686
Simplified flow chart for equivalent system computer program. ..................................... 691
FIGURE 255.
FIGURE 256.
FIGURE 257.
FIGURE 258.
FIGURE 259.
FIGURE 260.
FIGURE 261.
FIGURE 262.
FIGURE 263.
FIGURE 264.
FIGURE 265.
FIGURE 266.
FIGURE 267.
FIGURE 268.
FIGURE 269.
FIGURE 270.
FIGURE 271.
Supersedes pages iv - xiii of
MIL-STD-1797A
xxix
MIL-STD-1797A
APPENDIX A FIGURES
Figure
FIGURE 272.
FIGURE 273.
Page
Flow Chart for a modified Rosenbrock search algorithm................................................ 692
Example step-and-ramp HUD tracking sequences....................................................... 108n
FIGURE 274. Typical ATLAS light pattern (from NASA TM 101700) .................................... 108p
FIGURE 275 Adequate and desired performance for 4950th boom tracking
task (from 4950-FTR-93-05) ......................................................................................... 108q
FIGURE 276. Variation of crossover frequency with pitch attitude dynamics
(from AFFDL-TR-65-15) .................................................................................... 269
FIGURE 277. Specification of the criterion frequency.............................................................. 269
Supersedes pages iv - xiii of
MIL-STD-1797A
xxx
MIL-STD-1797A
TABLES
Table
TABLE I.
TABLE II.
TABLE III.
TABLE IV.
TABLE V.
TABLE VI.
Standard
Page
Operational Flight Envelope.............................................................................................. 21
Aircraft Normal States....................................................................................................... 24
Levels for Aircraft Failure States....................................................................................... 25
Pitch trim change conditions. ............................................................................................ 32
Flying qualities in atmospheric disturbances for Aircraft Normal States........................... 47
Flying qualities in atmospheric disturbances for Aircraft Failure States. .......................... 48
Supersedes pages iv - xiii of
MIL-STD-1797A
xxxi
MIL-STD-1797A
APPENDIX A TABLES
Table
TABLE I.
TABLE III.
TABLE V.
TABLE VI.
TABLE VII.
TABLE VIII.
TABLE IX.
TABLE X.
TABLE XI.
TABLE XII.
TABLE XIII.
TABLE XIV.
TABLE XV.
TABLE XVI.
TABLE XVII.
TABLE XVIII.
TABLE XIX.
TABLE XX.
TABLE XXI.
TABLE XXII.
Page
Operational Flight Envelope............................................................................................ 102
Levels for Aircraft Failure States..................................................................................... 119
Flying qualities in atmospheric disturbances for Aircraft Normal States......................... 655
Flying qualities in atmospheric disturbances for Aircraft Failure States. ........................ 655
Operational flight Envelope values. ................................................................................ 112
Recommended Levels for Aircraft Failure States. .......................................................... 120
Summary of minimum allowable intervention times for system failures. ........................ 154
Control-margin increments.............................................................................................. 163
Control surface lags. ....................................................................................................... 175
Examples of variations in LOES parameters with 1/Tθ2 fixed and free........................... 181
Lead/lag configurations with Level 1 LOES and Level 2 pilot ratings............................. 187
Classification of some known PIO cases (from NOR-64-143)........................................ 291
Pitch trim change conditions. .......................................................................................... 298
Maximum forces exerted on aircraft control stick (lb) by men and women
(AFAMRL-TR-81-39)....................................................................................................... 315
Pitch maneuvering force gradient limits. ......................................................................... 319
Category A control configurations for T-33 sidestick evaluations
(AFFDL-TR-79-3126). ..................................................................................................... 345
Experimental test points used in sidearm controller evaluations.................................... 346
Recommended pitch axis breakout forces (lb). .............................................................. 352
Recommended force limits for dives and recovery from dives. ...................................... 358
Guidance for lower limit on 1/Tθ2..................................................................................... 370
Conversion of ωsp Tθ2 to a phase angle criterion............................................................. 371
Recommended maximum roll-mode time constant (seconds). ...................................... 393
Spiral stability -- recommended minimum time to double amplitude. ............................. 408
Recommended minimum values for roll-spiral damping coefficient, ζRS ωRS .................. 412
Recommended allowable equivalent delay. ................................................................... 435
Roll performance for Class I and II aircraft. .................................................................... 451
Roll performance for Class III aircraft. ............................................................................ 451
Roll performance for Class IV aircraft. ............................................................................ 453
Flight phase CO roll performance in nominal 1-g rolls.................................................... 453
Flight phase CO roll performance, n > 1......................................................................... 454
Flight phase GA roll performance. .................................................................................. 454
YF-16 roll performance (from AFFTC-TR-75-15). .......................................................... 469
Maximum forces exerted on aircraft control stick (lb) by 61 men
and 61 women (AFAMRL-TR-81-39) .............................................................................. 488
TABLE XXXVI. Recommended maximum roll control force. ................................................................... 492
TABLE XXXVII. Recommended maximum roll control sensitivity for stickcontrolled aircraft............................................................................................................. 494
TABLE XXXIX. Recommended allowable breakout forces (lb). .............................................................. 512
TABLE XL.
Recommended minimum dutch roll frequency and damping. ........................................ 522
TABLE XLI.
Recommended bandwidth limits for wings-level turn mode. .......................................... 548
TABLE XLII.
Aircraft parameters subject to bandwidth limitation for wingslevel turn mode................................................................................................................ 548
TABLE XLIII. Recommended maximum sideslip excursions for large roll
control commands........................................................................................................... 557
'
(3) for r&o p& o < 0.07. ................................................................................. 570
TABLE XLIV. Limits on δrp
TABLE XXIII.
TABLE XXIV.
TABLE XXV.
TABLE XXVI.
TABLE XXVII.
TABLE XXVIII.
TABLE XXIX.
TABLE XXX.
TABLE XXXI.
TABLE XXXII.
TABLE XXXIII.
TABLE XXXIV.
TABLE XXXV.
TABLE XLV.
TABLE XLVI.
Ground rules for application of rating data to heading control criteria. ........................... 571
Parameters defining the LOS representation of the aileron-rudder
crossfeed......................................................................................................................... 576
Supersedes pages iv - xiii of
MIL-STD-1797A
xxxii
MIL-STD-1797A
APPENDIX A TABLES
Table
TABLE XLVII.
TABLE XLVIII.
TABLE XLIX.
TABLE L.
TABLE LI.
TABLE LI.
TABLE V.
TABLE VI.
TABLE LII.
TABLE LIII.
TABLE LIV.
TABLE LV.
TABLE LVI.
TABLE LVII.
Page
Physical interpretation of µ.............................................................................................. 578
Summary of current data. ............................................................................................... 581
Recommended minimum crosswind velocity requirements............................................ 587
Summary of Cooper-Harper pilot ratings for lateral translation mode. ........................... 616
Digest of pilot comments on specific aircraft high-AOA flying characteristics
(from AFWAL-TR-81-3108)............................................................................................. 629
Digest of pilot comments on specific aircraft high-AOA flying
characteristics (from AFWAL-TR-81-3108) .................................................................... 630
Flying qualities in atmospheric disturbances for Aircraft Normal States......................... 665
Flying qualities in atmospheric disturbances for Aircraft Failure States. ........................ 665
Recommended values for table V................................................................................... 666
Recommended values for table VI.................................................................................. 667
Digital filter implementation............................................................................................. 675
Examples of practical implemental matters. ................................................................... 688
A Survey of atmospheric disturbance models. ............................................................... 691
Linearized gust derivative terms in airframe dynamics................................................... 695
Supersedes pages iv - xiii of
MIL-STD-1797A
xxxiii
MIL-STD-1797A
APPENDIX A
AFFDL-TR-70-74
Vols I and II
An In-Flight Investigation to Develop Control System Design Criteria for
Fighter Airplanes; Neal, T. P. and Rogers E. Smith, December 1970
AFFDL-TR-70-145
An In-Flight Investigation of Lateral-Directional Dynamics for the Landing
Approach; Hall, G. W. and E. M. Boothe; October 1971
AFFDL-TR-70-155
Validation of the Flying Qualities Requirements of MIL-F-00878SA(USAF);
Brady, C. C. and J. Hodgkinson; January 1971
AFFDL-TR-71-134
Validation of the Flying Qualities Requirements of MIL-F-8785B(ASG),
Kandalaft, R. N.; September 1971
AFFDL-TR-71-164
In-Flight Investigation of an Unaugmented Class III Airplane in the Landing
Approach Task. Phase I: Lateral-Directional Study; Wasserman, R., F. F.
Eckhart and H. J. Ledder; January 1972
Vol I
AFFDL-TR-72-36
Evaluation of Lateral-Directional Handling Qualities and Roll-Sideslip
Coupling of Fighter Class Airplanes; Boothe, E. M. and M. L. Parrag, May
1972
AFFDL-TR-72-41
Revisions to MIL-F-8785B(ASG) Proposed by Cornell Aeronautical
Laboratory Under Contract F33615-71-C-1254; Chalk, C. R., D A. DiFranco,
et al.; April 1973
AFFDL-TR-72-141
Validation of the Flying Qualities Requirements of MIL-F-8785B(ASG) Using
the P-3B Airplane; Richards, R. B., D. L. Green and J. C. Rennie, November
1973
Vol I
AFFDL-TR-72-143
In-Flight Simulation of Minimum Longitudinal Stability for Large Delta-Wing
Transports in Landing Approach and Touchdown. Vol 1: Technical Results;
Wasserman, R. and J. F. Mitchell; February 1973
AFFDL-TR-73-76
Recommended Revisions to Selected Portions of MIL-F-8785B(ASG) and
Background Data; Ashkenas, I. L., R. H. Hoh and S. J. Craig; August 1973
AFFDL-TR-74-9
A Two-Phase Investigation of Longitudinal Flying Qualities for Fighters;
Boothe, E. M., R. T. N. Chen and C. R. Chalk; April 1974
AFFDL-TR-74-61
Investigation of Flying Qualities of Military Aircraft at High Angles of Attack.
Vol 1: Technical Results; Johnston, D. E., I. L. Ashkenas and J. R. Hogge;
June 1974
AFFDL-TR-74-130
(2 Vols)
Extension of the Method for Predicting Six-Degree-of-Freedom Store
Separation Trajectories at Speeds Up to the Critical Speed to Include A
Fuselage with Noncircular Cross Section; Dillenius, M. F. E., F. K. Goodwin
and J. N. Nielsen; November 1974
AFFDL-TR-75-3
Evaluation of the Flying Qualities Requirements of MIL-F-8785B(ASG) Using
the C-5A Airplane-, Silvers, C. L. and C. C. Withers; March 1975
AFFDL-TR-76-78
Direct Side Force Control Criteria for Dive Bombing. Vol 1: Summary Vol II:
Analysis and Results; Brulle, R. V., W. A. Moran and R. G Marsh;
September 1976
REPRINTED WITHOUT CHANGE
61
MIL-STD-1797A
APPENDIX A
AFFDL-TR-77-34
Limited-Flight Evaluation of Sidestick Controller Force-Deflection
Characteristics on Aircraft Handling Qualities; Cima, William M., et al.; Nov
1977
AFFDL-TR-77-57
A Theory for Longitudinal Short-Period Pilot Induced Oscillations; Smith,
Ralph H.; June 1977
AFFDL-TR-78-9
Fighter CCV Phase IV Report, Vol II: Flight Test Data Evaluation. Vol III:
Test Phase Data Summary, Parts I and 2; McAllister, J. D., et al.; February
1978
AFFDL-TR-78-122
Effects of Control System Dynamics on Fighter Approach and Landing
Longitudinal Flying Qualities (Volume I); Smith, Rogers E.; March 1978
AFFDL-TR-78-154
Handling Quality Requirements for Advanced Aircraft Design: Longitudinal
Mode; Smith, Ralph H. and Norman D. Geddes; Aug 1979
AFFDL-TR-78-171
Proceedings of AFFDL Flying Qualities Symposium Held at Wright State
University 12-15 September, 1978; Black, G. T., Moorhouse, D. J., et al.,
compilers; December 1978:
“Task-Oriented Flying Qualities for Air-to-Ground Gun Attack;" Brandeau, G.
“B-1 Experience Related to MIL-F-8785B and Proposed Revisions;"
Campbell, J. E.
"An Approach to Simplify the Specification of Low-Speed Maneuvering Pitch
Control Force;" Cichy, D. R.
"High Angle of Attack Flying Qualities and Departure Criteria Development;"
Hellman, G. K. and R. B. Crombie
"Northrop Review of MIL-F-8785B Proposed Revision;" Lockenour, J
“Evaluation of Selected Class III Requirements of MIL-F-8785B(ASG),
‘Flying Qualities of Piloted Airplanes;’” Withers, C. C.
“Flying Qualities of Piloted Airplanes;'" Withers, C. C.
"Discussion and Status of the Proposed Revision (1978) to MIL-F-8785B;"
Moorhouse, D. J., R. J. Woodcock and T. P. Sweeney
AFFDL-TR-79-3126
Flight Qualities Design Requirements for Sidestick Controllers; Black, G. T.
and D. J. Moorhouse; October 1979
Supersedes pages 62-75 of
MIL-STD-1797A
62
MIL-STD-1797A
APPENDIX A
AFWAL-TR-80-3032
Prediction of Supersonic Store Separation Characteristics Including
Fuselage and Stores of Noncircular Cross Section (4 volumes); Goodwin, F.
K., M. F. E. Dillenius and J. Mullen. Jr.; November 1980
AFWAL-TR-80-3060
Simulation Analysis: Unorthodox Control Force Fighter Aircraft, Vol II:
Detailed Summary; Mitchell, A. L., et al.; April 1980
AFWAL-TR-80-3067
Flying Qualities Design Criteria: Proceedings of AFFDL Flying Qualities
Symposium Held at Wright-Patterson Air Force Base in October 1979.
Crombie, R. B. and D. J. Moorhouse, compilers; May 1980
AFWAL-TR-80-3141
Investigation of High-Angle-of-Attack Maneuvering - Limiting Factors, Part 1:
Analysis and Simulation; Johnston, D. E., D. G. Mitchell and T T. Myers;
December 1980
AFWAL-TR-81-3027
Development of Handling Quality Criteria for Aircraft with Independent
Control of Six-Degrees-of-Freedom; Hoh, R. H., T. T. Myers, et al.; April
1981
AFWAL-TR-81-3090
Notes on Lateral-Directional Pilot Induced Oscillations; Sith, Ralph H.; Mar
1982
AFWAL-TR-81-3108
Investigation of High AOA Flying Qualities and Design Guides; Johnston, D.
E. and R. K. Heffley; December 1981
AFWAL-TR-81-3109
Background Information and User Guide for MIL-F-8785C.
Military
Specification - Flying Qualities of Piloted Airplanes; Moorhouse, D J. and R.
J. Woodcock; September 1981
AFWAL-TR-81-3116
Equivalent System Verification and Evaluation of Augmentation Effects on
Fighter Approach and Landing Flying Qualities; Smith, Rogers E.;
September 1981
AFWAL-TR-81-3118
In-Flight Investigation of Large Airplane Flying Qualities for Approach and
Landing; Weingarten, N. C. and C. R. Chalk; September 1981
AFWAL-TR-81-3171
Lateral Flying Qualities of Highly Augmented Fighter Aircraft, Vols. I and 11;
Monegan, S. J., Rogers E. Smith and R. E. Bailey; June 1982
AFWAL-TR-82-3014
Proposed Revisions to MIL-F-8785C Related to Flight Safety of Augmented
Aircraft, 3 Vols.; Schuler, J. M. and M A. Dahl, April 1982
AFWAL-TR-82-3064
Design Criteria for the Future of Flight Controls, Proceedings of the Flight
Dynamics Laboratory Flying Qualities and Flight Control Symposium, 2-5
March, 1982; Fuller, S. G. and Potts, D. W., compilers; July 1982
AFWAL-TR-82-3081
Proposed MIL Standard and Handbook - Flying Qualities of Air Vehicles, Vol
II: Proposed MIL Handbook; Hoh, R. H., Mitchell, D. G., et al November
1982
Supersedes pages 62-75 of
MIL-STD-1797A
63
MIL-STD-1797A
APPENDIX A
AFWAL-TR-83-3015
Suggested Revisions to MIL-F-8785C for Large Class III Aircraft; Nleyer,
R. T., et al.; February 1983
AFWAL-TR-86-3093
Proceedings of the Twenty-Second Annual Conference on Manual
Control, Dec 1986
“Test Pilot Evaluation of the Closed-Loop GRATE Flight Test Technique;”
Biezad, Daniel J. and Steven R. Sturmer
AFWAL-TM-87-180-FIGC
S/MTD Series IV Manned Simulation: An Air Force Evaluation of S/MTD
Flaps Down Operation; Leggett, David B., et al.; Aug 1987
AFFDL-FGC-TM-71-7
Validation of the Handing Qualities Degradation Probabilities of MIL-F008785A Using F-4C Air Force Manual 66-1 Maintenance Data, Ullman,
Lt., T. Calanducci, and Lt. Linck; August 1971
AFAMRL-TR-73-78
Manual Control Performance and Dynamic Response During Sinusoidal
Vibration; Allen, R. Wade, Henry R. Jex, and Raymond E. Magdaleno,
October 1973
AFAMRL-TR-81-39
Male and Female Strength Capabilities for Operating Aircraft Controls;
McDaniel, Joe W.; March 1981
AFFTC-SD-69-5
A-7D Stability and Control Military Preliminary Evaluations (Phase IA and
IB); Gobert, Don 0. and William T. Twinting; April 1969
AFFTC-TD-75-1
Tracking Test Techniques for Handling Qualities Evaluation; Twisdale, T
R. and D. L. Franklin; May 1975
AFFTC-TR-75-15
Flying Qualities Evaluation of the YF-16 Prototype Lightweight Fighter,
Eggers, James A. and William F. Bryant, Jr.; July 1975
AFFTC-TR-75-32
F-15A Approach-to-Stall/Stall/Post-Stall Evaluation; Wilson, Donald B and
Charles A. Winters; January 1976
AFFTC-TR-76-15
Flight Test Development and Evaluation of a Multimode Digital Flight
Control System Implemented in an A-7D (DIGITAC), Damman, Lawrence,
Robert Kennington, Paul Kirsten, Ronald Grabe, and Patrick Long; June
1976
AFFTC-TR-77-11
A-10A Flying Qualities Air Force Developmental Test and Evaluation;
Stewart, Will R., et al; Jul 1975
AFFTC-TR-77-23
YF-16 Control Configured Vehicle (CCV) Operational Potential, Flying
Qualities, and Performance Evaluation; Wood, Richard A., et al; Jan 1978
AFFTC-TR-77-27
System Identification from Tracking (SIFT), a New Technique for Handling
Qualities Test and Evaluation (Initial Report); Twisdale, T. R and T. A.
Ashurst; November 1977
AFFTC-TR-79-2
Flying Qualities and Flight Control System Evaluation of the B-1 Strategic
Bomber; Ross, Jerry L., Page G. McGirr, and Otto J. Waniczek, Jr.; May
1979
AFFTC-TR-79-10
F-16A/B Flying Qualities Full-Scale Development Test and Evaluation;
Pape, James A. and Michael P. Garland; September 1979
Supersedes pages 62-75 of
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64
MIL-STD-1797A
APPENDIX A
AFFTC-TR-79-18
F-16A/B High Angle of Attack Evaluation; Wilson, Donald B. and Robert C.
Ettinger; October 1979
AFFTC-TR-80-23
F-15C Flying Qualities Air Force Development Test and Evaluation;
Shaner, Keith L. and Robert W. Barham; November 1980
AFFTC-TR-80-29
F-16 Flying Qualities with External Stores; Garland, Michael P., Michael K.
Nelson, and Richard C. Patterson; February 1981
AFFTC-TR-83-45
AFTI/F-16 Handling Qualities Evaluation; Sorokowski, Paul J., et al; Feb
1984
AFFTC-TR-91-29
STOL/Maneuver Technology Demonstrator Flying Qualities and Integrated
Flight/Propulsion Control System Evaluation; Crawford, Mark R. and
Michael J. Costigan; Dec 1991
AFFTC-TLR-93-41
Human Pilot Response During Single- and Multi-Axis Tracking Tasks;
Edkins, Craig R.; Dec 1993
FDL-TDR-64-60
Flight Evaluation of Various Short Period Dynamics at Four Drag
Configurations for the Landing Approach Task; Chalk, C. R., October
1964; Chalk, Charles R.; October 1964
FTC-TR-66-24
Frequency Response Method of Determining Aircraft Longitudinal Short
Period Stability and Control System Characteristics in Flight; Klung, H A.,
Jr.; August 1966
FTC-TR-67-19
Evaluation of Longitudinal Control Feel System Modifications Proposed for
USAF F/RF-4 Aircraft ; Keith, L. A., R. R. Richard, and G J. Marrett,
December 1968
FTC-TD-72-1
Development and Evaluation of the TWeaD II Flight Control Augmentation
System; Carleton, David L., Richard E. Lawyer, and Cecil W. Powell;
November 1972
FTC-TD-73-2
Background Information and User Guide for MIL-S-83691; Sharp, Patrick
S. and Collet E. McElroy, March 1974
FTC-TR-73-32
Air Force Evaluation of the Fly-by-Wire Portion of the Survivable Flight
Control System Advanced Development Program; Majoros, Robert L.;
August 1973
FTC-TIH-79-2
USAF Test Pilot School, Flying Qualities Handbook, Flying Qualities
Theory and Flight Test Techniques; November 1979
USAFTPS Report 82B-4
Limited Comparison of Longitudinal Flying qualities Between the SAFTD
Ground Simulator and the NT-33A In-flight Simulator; Payne, James M., et
al; 31 May 1983
USAFTPS-TR-88A-TM1
Adaptable Target Lighting Array System (HAVE ATLAS); Wilson, E. M., et
al; Dec 1988
USNTPS-FTM-103
Fixed Wing Stability and Control, Theory and Flight Techniques; 1
November 1981
WL-TR-92-3027
Lessons Learned from the STOL and
Demonstrator; Moorhouse, David J.; Jun 1993
Supersedes pages 62-75 of
MIL-STD-1797A
65
Maneuver
Technology
MIL-STD-1797A
APPENDIX A
WL-TR-93-3081
Aircraft Maneuvers for the Evaluation of Flying Qualities and Agility, Vol 1:
Maneuver Development Process and Initial Maneuver Set; Wilson, David
J., David R. Riley, and Kevin D. Citurs; McDonnell Douglas Aerospace;
Aug 1993
WL-TR-93-3082
Aircraft Maneuvers for the Evaluation of Flying Qualities and Agility, Vol 2:
Maneuver Descriptions and Selection Guide; Wilson, David J., David R.
Riley, and Kevin D. Citurs; McDonnell Douglas Aerospace; Aug 1993
WL-TR-93-3083
Aircraft Maneuvers for the Evaluation of Flying Qualities and Agility, Vol 3:
Simulation Data; Wilson, David J., David R. Riley, and Kevin D. Citurs;
McDonnell Douglas Aerospace; Aug 1993
WRDC-TR-89-3036
Results of a Piloted Simulation of the STOL/Maneuver Technology
Demonstrator; McDonnell Douglas Aerospace; Aug 1993
4950-FTR-93-05
Closed Loop Handling Tasks Testing and Verification for Class III-L
Aircraft; DeWitt, Bruce R.; Oct 1993
FAA FAR Part 23
Airworthiness Standards:
Airplanes; June 1974
FAA FAR Part 25
Airworthiness Standards: Transport Category Airplanes; June 1974
FAA-ADS-69-13
An In-Flight Investigation of Lateral-Directional Dynamics for Cruising
Flight; Hall, G. W.; December 1969
FAA-RD-70-61
A Flight Simulator Study of STOL Transport Lateral Control
Characteristics; Drake, Douglas E., Robert A. Berg, Gary L. Teper, and W.
Allen Shirley; September 1970
FAA-RD-70-65
Flying Qualities of Small General Aviation Airplanes. Part 2: The
Influence of Roll Control Sensitivity Roll Damping, Dutch-Roll Excitation,
and Spiral Stability; Ellis, David R; April 1970
FAA-RD-74-206
Wind Models for Flight Simulator Certification of Landing and Approach
Guidance and Control Systems; Barr, Neal M., Dagfinn Gangsaas, and
Dwight R. Schaeffer; December 1974
FAA-RD-75-123
Identification of Minimum Acceptable Characteristics for Manual STOL
Flight Path Control; Hoh, Roger H., Samuel J. Craig, and Irving L.
Ashkenas; June 1976
FAA-RD-77-25
A Study of Lightplane Stall Avoidance and Suppression; Ellis, David R.;
February 1977
FAA-RD-77-36
Wind Shear Modeling for Aircraft Hazard Definition; Frost, Walter and
Dennis W. Camp; March 1977
FAA-RD-77-173
Proceedings of the First Annual Meteorological and Environmental Inputs
to Aviation Systems Workshop. "Wind Models for Flight Simulator
Certification of Landing and Approach Guidance and Control Systems",
Schaeffer, Dwight R.; March 1977
Supersedes pages 62-75 of
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66
Normal, Utility, and Acrobatic Category
MIL-STD-1797A
APPENDIX A
FAA-RD-78-7
Simulation and Analysis of Wind Shear Hazard; Lehman, John Ni Robert
K. Heffley, and Warren F. Clement; December 1977
FAA-RD-79-59
Powered-Lift Aircraft Handling Qualities in the Presence of NaturallyOccurring and Computer-Generated Atmospheric Disturbances; Jewell,
Wayne F., Warren F. Clement, Thomas C. West, and S. R. Ni Sinclair;
May 1979
FAA-RD-79-84
Piloted Flight Simulation Study of Low-Level Wind Shear, Phase 4; Foy,
W. H. and W. B. Gartner; March 1979
FAA Advisory
High-Speed Characteristics; 24 November 1965
Circular AC25.253-1A
DOT/FAA/CT-82/130-II
Flying Qualities of Relaxed Static Stability Aircraft, Vol II; McRuer, D. T
and T. T. Myers; September 1982
NACA Memo Rpt L6E20
Flight Investigation to Improve the Dynamic Longitudinal Stability and
Control-Feel Characteristics of the P-63A-1 Airplane with Closely
Balanced Experimental Elevators, Johnson, Harold I.; July 1946
NASA Memo 1-29-59A
A Pilot Opinion Study of Lateral Control Requirements for Fighter- Type
Aircraft; Creer, Brent Y., John D. Stewart, Robert B. Merrick, and Fred J.
Drinkwater III; March 1959
NASA Memo 12-10-58A
A Flight Investigation to Determine the Lateral Oscillatory Damping
Acceptable for an Airplane in the Landing Approach; McNeill, Walter E.
and Richard F. Vomaske-, February 1959
NASA-CP-2028
Proceedings of the First Annual Meteorological and Environmental Inputs
to Aviation Systems Workshop, 'Wind Models for Flight Simulator
Certification of Landing and Approach Guidance and Control Systems";
Schaeffer, Dwight R.; March 1977
NASA-CP-2428
Twenty-First Annual Conference on Manual Control; “A Flight Test Method
for Pilot/Aircraft Analaysis;” Koehler, Ruthard and Ernst Buchacker;
DFVLR; May 1986
NASA-CR-239
Development of Satisfactory Lateral-Directional Handling Qualities in the
Landing Approach; Stapieford, Robert L., Donald E. Johnston, Gary L.
Teper, and David H. Weir; July 1965
NASA-CR-635
In-Flight and Ground Based Simulation of Handling Qualities of Very Large
Airplanes in Landing Approach; Condit, Philip M., Laddie G. Kimbrel, and
Robert G. Root; October 1966
NASA-CR-778
Evaluation of Lateral-Directional Handling Qualities of Piloted Re-EntryVehicles Utilizing Fixed-Base and In-Flight Evaluations; Meeker, J I May
1967
NASA-CR-2017
Handling Qualities Criteria for the Space Shuttle Orbiter During the
Terminal Phase of Flight; Stapleford, Robert L., Richard H. Klein, and
Roger H. Hoh; April 1972
Supersedes pages 62-75 of
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MIL-STD-1797A
APPENDIX A
NASA-CR-2451
Non-Gaussian Atmospheric Turbulence Model for Use in Flight
Simulators; Reeves, P. M., G. S. Campbell, V. M. Ganzer, and R. G.
Joppa; September 1974
NASA-CR-2677
Manual and Automatic Flight Control During Severe Turbulence
Penetration; Johnston, Donald E., Richard H. Klein, and Roger H. Hoh,
April 1976
NASA-CR-152064
Investigation of the Vulnerability of Powered Lift STOLs to Wind Shear;
Hoh, Roger H. and Wayne F. Jewell; October 1976
NASA-CR-152139
Study of a Safety Margin System for Powered-Lift STOL Aircraft; Heffley,
Robert K. and Wayne F. Jewell; May 1978
NASA-CR-152194
A Study of Key Features of the RAE Atmospheric Turbulence Model;
Jewell, Wayne F. and Robert K. Heffley; October 1978
NASA-CR-159059
An Investigation of Low-Speed Lateral Acceleration Characteristics of
Supersonic Cruise Transports Using the Total In-Flight Simulator (TIFS),
Weingarten, N. C.; July 1979
NASA-CR-159236
Calspan Recommendations for SCR Flying Qualities Design Criteria-,
Chalk, C. R.; April 1980
NASA-CR-163108
Analyses of Shuttle Orbiter Approach and Landing Conditions; Teper,
Gary L., Richard J. DiMarco, Irving L. Ashkenas, and Roger H. Hoh; July
1981
NASA-CR-172491
Pitch Rate Flight Control Systems in the Flared Landing Task and Design
Criteria Development; Berthe, C. J., C. R. Chalk, and S. Sarrafian
NASA-CR-177331
Mission-Oriented Requirements for Updating MIL-H-8501, Vols. I and II;
Clement, W. F., et al.; January 1985
NASA-CR-178188
Flared Landing Approach Flying Qualities; Weingarten, Norman C., et al;
Dec 1986
NASA-TM-86728
Application of Frequency Domain Handling Qualities Criteria to the
Longitudinal Landing Task; Sarrafian, S. K. and B. G. Powers; August
1985
NASA-TM-101700
Initial Flight Test of a Ground Deployed System for Flying Qualities
Assessment; Shafer, M. F., et al; Aug 1989
NASA-TM-X-62
Motion Simulator Study of Longitudinal Stability Requirements for Large
Delta Wing Transport Airplanes During Approach and Landing with
Stability Augmentation Systems Failed; Snyder, C. T., E. B. Fry, et al.,
December 1972
NASA-TM-X-1584
A Review of Transport Handling-Qualities Criteria in Terms of Preliminary,
XB-70 Flight Experience; Powers, Bruce G.; May 1968
NASA-TN-D-173
Flight Investigation of Automatic Stabilization of an Airplane Having Static
Longitudinal Instability; Russell, Walter R., S. A. Sjoberg, and William L
Alford; December 1959
Supersedes pages 62-75 of
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68
MIL-STD-1797A
APPENDIX A
NASA-TN-D-211
Flight Investigation of Pilot's Ability to Control an Airplane Having Positive
and Negative Static Longitudinal Stability Coupled with Various Effective
Lift-Curve Slopes; Brissenden, Roy F., William L Alford, and Donald L.
Mallick; February 1960
NASA-TN-D-746
Flight Controllability Limits and Related Human Transfer Functions as
Determined from Simulator and Flight Tests, Taylor, Lawrence W. and
Richard E. Day; May 1961
NASA-TN-D-779
Flight Investigation Using Variable-Stability Airplanes of Minimum Stability,
Requirements for High-Speed, High-Altitude Vehicles; McFadden, Norman
M., Richard F. Vomaske, and Donovan R Heinle; April 1961
NASA-TN-D-792
Attitude Control Requirements for Hovering Control Through the Use of a
Piloted Flight Simulator; Faye, A. E., Jr.; April 1961
NASA-TN-D-1141
The Effect of Lateral-Directional Control Coupling on Pilot Control of an
Airplane as Determined in Flight and in a Fixed-Base Flight Simulator;
Vomaske, Richard F., Melvin Sadoff, and Fred J. Drinkwater III; November
1961
NASA-TN-D-1328
A Flight Determination of the Attitude Control Power and Damping
Requirements for a Visual Hovering Task in the Variable Stability and
Control X-14A Research Vehicle; Rolls, L. S. and F. J. Drinkwater; May
1962
NASA-TN-D-1552
A Study of a Pilot's Ability to Control During Simulated Augmentation
System Failures; Sadoff, Melvin; November 1962
NASA-TN-D-1888
A Preliminary Study of Handling-Qualities Requirements of Supersonic
Transports in High-Speed Cruising Flight Using Piloted Simulators; White,
Maurice D., Richard F. Vomaske, Walter E. McNeill, and George E
Cooper; May 1963
NASA-TN-D-2251
A Piloted Simulator Study of Longitudinal Handling Qualities of Supersonic
Transport in the Landing Maneuver; Bray, Richard S., April 1964
NASA-TN-D-3726
An Evaluation of the Handling Qualities of Seven General-Aviation Aircraft;
Barber, Marvin R., Charles K. Jones, Thomas R. Sisk, and Fred W. Haise;
November 1966
NASA-TN-D-3910
A Simulator and Flight Study of Yaw Coupling in Turning Maneuvers of
Large Transport Aircraft; McNeill, W. E. and R. C. Innis; May 1967
NASA-TN-D-3971
Determination of Flight Characteristics of Supersonic Transports During
the Landing Approach with a Large Jet Transport In-Flight Simulator; June
1967
NASA-TN-D-5153
The Use of Pilot Rating in the Evaluation of Aircraft Handling Qualities;
Cooper, G. E. and Harper, R. P., Jr.; April 1969
Supersedes pages 62-75 of
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MIL-STD-1797A
APPENDIX A
NASA-TN-D-5466
Simulator Study of Coupled Roll-Spiral Mode Effects on Lateral-Directional
Handling Qualities; Grantham, W. D., F. L. Moore, P. L. Deal, and J. M.
Patton, Jr.; March 1970
NASA-TN-D-5957
Flight Investigation of the Roll Requirements for Transport Airplanes in
Cruising Flight; Holleman, Euclid C.; September 1970
NASA-TN-D-6496
Analysis of a Coupled Roll-Spiral-Mode, Pilot-Induced Oscillation
Experienced with the M2-F2 Lifting Body; Kempel, R. W.; September 1971
NASA-TN-D-6811
In-Flight Pilot Evaluations of the Flying Qualities of a Four-Engine Jet
Transport; Holleman, Euclid C. and Glenn B. Gilyard; May 1972
NASA-TN-D-7703
Flight Investigation of Advanced Control Systems and Displays for a
General Aviation Airplane; Loschke, Paul C., Marvin R. Barber, Einar K
Enevoldson, and Thomas C. McMurtry; June 1974
NASA-TP-1368
Flight Comparison of the Transonic Agility of the F-111A Airplane and the
F-111 Supercritical Wing Airplane; Friend, Edward L. and Glenn M.
Sakamoto; December 1978
NASA-TR-R-199
Dynamic Response of Airplanes to Atmospheric Turbulence Including
Flight Data on Input and Response; Houbolt, John C., Roy Steiner, and
Kermit G. Pratt; June 1964
NADC-ED-6282
Proposal for a Revised Military Specification, 'Flying Qualities of Piloted
Airplanes' (MIL-F-8785ASG) with Substantiating Text; Mazza, C. J.,
William Becker, et al.; 22 July 1963
NADC-76154-30
"Design Charts and Boundaries for Identifying Departure Resistant Fighter
Configurations;" Bihrie, W., Jr. and Barnhart, B.; July 1978
NADC-77052-30
Development of VTOL Flying Qualities Criteria for Low Speed and Hover;
Hoh, Roger H. and Irving L. Ashkenas; December 1979
NADC-78182-60
Development and Analysis of a CVA and a 1052 Class Fast Frigate Air
Wake Model; Nave, Ronald L.; September 1978
NADC-81186-60
The Control Anticipation Parameter for Augmented Aircraft, Bisclioff, D E.;
May 1981
NADC-85091-60
“Investigation of Departure Susceptibility Criteria Using the Dynamic Flight
Simulator;" Rhodeside, G.; June 1985
NADC-85130-60
Guidelines for Designing Flying qualities Experiments; Dukes, Theodor A.;
Jun 1985
20.4 Nongovernment documents
AIAA Paper 64-353
Jet Transport Operation in Turbulence; Soderlind, Paul A.; July 1964
AIAA Paper 69-898
Summary and Interpretation of Recent Longitudinal Flying Qualities
Results; Ashkenas, I. L.; August 1969
AIAA Paper 75-985
In-Flight Simulation-of the Light Weight Fighters; Hall, G. W. and R. P.
Harper; August 1975
Supersedes pages 62-75 of
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MIL-STD-1797A
APPENDIX A
AIAA Paper 77-1119
Direct-Force Flight-Path Control--the New Way to Fly; Watson, John H.
and Jack D. McAllister; August 1977
AIAA Paper 77-1122
Equivalent System Approaches to Handling Qualities Analysis and Design
Problems in Augmented Aircraft; Hodgkinson, J. and W. J. LaManna, 8-10
August 1977
AIAA Paper 77-1145
A Study of Key Features of Random Atmospheric Disturbance Models for
the Approach Flight Phase; Heffley, Robert K.; August 1977
AIAA Paper 78-1500
Rolling Tail Design and Behavior as Affected by Actuator Hinge Moment,
Ball, J. M.; August 1978
AIAA Paper 79-1783
Initial Results of an Inflight, Simulation of Augmented Dynamics in Fighter
Approach and Landing; Hodgkinson, J. and K. A. Johnston; 6-8 August
1979
AIAA Paper 79-1962
Flight Tests of a Microprocessor Control System; Stengel, R F. and G. E.
Miller; October 1979
AIAA Paper 80-0703
Review of Nonstationary Gust-Responses of Flight Vehicles, Gaonkar, G.
H.; July 1980
AIAA Paper 80-1611-CP
Flight Evaluation of Augmented Fighter Aircraft; Hodgkinson, J. and R. C.
Snyder; 11-13 August 1980
AIAA Paper 80-1626-CP
A Summary of an In-Flight Evaluation of Control System Pure Time Delays
During Landing Using the F-8 DFBW Airplane; Berry, D. T., B G. Powers,
K. J. Szalai, and R. J. Wilson; 11-13 August 1980
AIAA Paper 80-1627-CP
Low Order Equivalent Models of Highly Augmented Aircraft Determined
from Flight Data Using Maximum Likelihood Estimation; Shafer, M. F; 1113 August 1980
AIAA Paper 80-1628-CP
Handling Qualities Criteria for Wing-Level-Turn Maneuvering During an Air
to Ground Delivery; Sammonds, R. I. and J. W. Bunnell, Jr.; August 1980
AIAA Paper 80-1633
Identification of Flexible Aircraft from Flight Data; Eulrick, B. J. and E. D.
Rynaski; August 1980
AIAA Paper 80-1836
The Turbulent Wind and Its Effect on Flight; Etkin, B.; August 1980
AIAA Paper 81-0302
Atmospheric Disturbance Models and Requirements for the Flying
Qualities Military Standard and Handbook; Heffley, R. K., W. F. Jewell, R.
H. Hoh, and D. J. Moorhouse; January 1981
AIAA Paper 87-2561
Analysis and Application of Aircraft Departure Prediction Criteria to the
AV-8B Harrier II; Tinger, H.L.; August 1987
AIAA Paper 89-3358
Cooper-Harper Pilot Rating Variability; Wilson, David J. and David R.
Riley; Aug 1989
AIAA Paper 90-2822
More on Cooper-Harper Pilot Rating Variability; Riley, David R. and David
J. Wilson; Aug 1990
AIAA Paper 93-3645
Development of Flying Qualities and Agility Evaluation Maneuvers; Wilson,
David J., David R. Riley, Kevin D. Citurs, and Thomas J. Cord; AIAA
Atmospheric Flight Mechanics Conference; Monterey CA; 9-11 Aug 1993
Supersedes pages 62-75 of
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MIL-STD-1797A
APPENDIX A
AIAA Paper 93-3816
Initial Results of an In-Flight Investigation of Longitudinal Flying Qualities
for Augmented large Transports in Approach and Landing; Rossitto, K., et
al; Aug 1993
SAE ARP 842B
Design Objectives for Flying Qualities of Civil Transport Aircraft
Delft Univ of Tech Memo
M-304
Non-Gaussian Structure of the Simulated Turbulent Environment in Piloted
Flight Simulation; van de Moeskijk, G. A. J.; April 1978
Princeton Univ Rpt 604
A Study of Pilot-Induced Lateral-Directional Instabilities; Caporali, R. L., J.
P. Lamers, and J. R. Totten; May 1962
Princeton Univ Rpt 727
Lateral-Directional Flying Qualities for Power Approach; Seckel, E., G. E.
Miller, and W. B. Nixon; September 1966
Princeton Univ Rpt 777
Comparative Flight Evaluation of Longitudinal Handling Qualities in Carrier
Approach; Eney, J. A.; May 1966
Princeton Univ Rpt 797
Lateral-Directional Flying Qualities for Power Approach: Influence of Dutch
Roll Frequency; Seckel, E., J. A. Franklin, and G E. Miller, September
1967
Stanford Univ
Wind Modeling and Lateral Aircraft Control for Automatic Landing; Holley,
William E. and Arthur E. Bryson; January 1975
SUDAAR No. 489
ARC R&M No. 917
Preliminary Report on Behavior of Aeroplanes When Flying Inverted with
Special Reference to Some Accidents on "A";O'Gorman, Mervyn,
Chairman, Accidents Committee; January 1919
ESDU Item No. 74031
Characteristics of Atmospheric Turbulence Near the Ground. Part II:
Single Point Data for Strong Winds (Neutral Atmosphere), October 1974
ESDU Item No. 75001
Characteristics of Atmospheric Turbulence Near the Ground. Part III
Variations in Space and Time for Strong Winds (Neutral Atmosphere), July
1975
IAS Paper 60-18
Development of Lateral-Directional Flying Qualities Criteria for Supersonic
Vehicles, Based on a Stationary Flight Simulator Study; Crone, R M. and
R. C. A'Harrah; January 1960
ICAS-86-5.3.4
Handling Qualities for Unstable Combat Aircraft; Gibson, J. C.; September
1986
MDC Rpt A5596
Flying Qualities Analysis of an In-Flight Simulation of High Order Control
System Effects on Fighter Aircraft Approach and Landing-, Johnston, K A.
and J. Hodgkinson, 22 December 1978
MDC Rpt A6792
Definition of Acceptable Levels of Mismatch for Equivalent Systems of
Augmented Aircraft; Wood, J. R. and J. Hodgkinson; 19 December 1980
NLR-TR-79127U
Determination of Low-Speed Longitudinal Maneuvering Criteria for
Transport Aircraft with Advanced Flight Control Systems; Mooij, H. A., W.
P. Boer, and M. F. C. van Gool; 1979
NLR Memorandum
A Digital Turbulence Model for the NLR Moving - Base Flight Simulator,
Part I; Jansen, C. J., August 1977
VS-77-024
Supersedes pages 62-75 of
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72
MIL-STD-1797A
APPENDIX A
NLR Memorandum
VS-77-025
A Digital Turbulence Model for the NLR Moving - Base Flight Simulator,
Part II; Jansen, C. J.; August 1977
Boeing D6-10725
A Simulator and Flight Evaluation of the Longitudinal and Lateral Control
Requirements of the C-5A for the Landing Approach Task; Eldridge, W 18
May 1965
Boeing D6-10732 T/N
A Note on Longitudinal Control Response; Higgins, H. C.; June 1965
Calspan Rpt No. 7738-24
NT-33A In-flight Investigation into Flight Control System Rate Limiting;
Ohmit, E. E.; Feb 1994
Calspan FRM No. 554
The Ideal Controlled Element for Real Airplanes Is Not K/s; Chalk, C. R.;
August 1981
Comell Aero Lab
Flight Evaluation of a Stability Augmentation System for Light Airplanes;
Eckhart, F. F., G. W. Hall, and P. A. Martino; November 1966
IH-2154-F-1
Cornell Aero Lab
TB-574-F-3
Cornell Aero Lab
TB-574-F-6
Cornell Aero Lab
TB-1094-F-I
Cornell Aero Lab
TB-1444-F-I
Douglas Aircraft Co.
LB-25452
General Dynamics Rpt
A Flight Investigation of Minimum Acceptable Lateral Dynamic Stability,
Graham, D. and C. James; 30 April 1950
A Flight Investigation of Acceptable Roll to Yaw Ratio of the Dutch Roll
and Acceptable Spiral Divergence; Bull, G.; February 1952
Flight Evaluations of the Effect of Variable Spiral Damping in a JTB-26B
Airplane; Rhoads, D. W.; October 1957
Handling Qualities Requirements as Influenced by Prior Evaluation Time
and Sample Size; Kidd, E. A. and G. Bull; February 1963
Investigation of Pilot-Induced Longitudinal Oscillation in the Douglas
Model A4D-2 Airplane; Terrill, W. H., J. G. Wong, and L. R. Springer; 15
May 1959
9 December 1968
FZM-12-2652
Norair Rpt No.
NOR-64-143
Systems Tech. Inc.
TR-124-1
Systems Tech. Inc.
Pilot Induced Oscillations: Their Cause and Analysis; Ashkenas, Irving L.,
Henry R. Jex, and Duane T. McRuer; June 1964
A Systems Analysis of Longitudinal Piloted Control in Carrier Approach,
Cromwell, C. J. and I. L. Ashkenas; June 1962
Carrier Landing Analyses; Durand, Tulvio; February 1967
TR-137-2
Systems Tech. Inc.
TR-189-1
Systems Tech. Inc.
TR-190-1
Systems Tech. Inc.
TR-199-1
Supersedes pages 62-75 of
MIL-STD-1797A
Background Data and Recommended Revisions for MIL-F-8785B(ASG),
'Military Specification -- Flying Qualities of Piloted Airplanes'; Craig,
Samuel J. and Irving L. Ashkenas; March 1971
Outsmarting
MIL-F-8785B(ASG),
the
Military
Flying
Qualities
Specification; Stapleford, Robert L., Duane T. McRuer, Roger H. Hoh, et
al; August 1971
Analytical Assessment of the F-14 Aircraft Control and Handling
Characteristics; Johnston, Donald E. and Samuel J. Craig; February 1972
73
MIL-STD-1797A
APPENDIX A
Systems Tech. Inc.
TR-1090-1
Systems Tech. Inc.
WP-189-3
Vought Corp Rpt No.
2-55800/8R-3500
Analytical Assessment of the F-18A Flying Qualities During Carrier
Approach; Ringland, R. F. and D. E. Johnston; September 1977
Effect of Sideslip on Precise Lateral Tracking; Hoh, R. H. and H. R. Jex;
November 1969
Mathematical Models for the Aircraft Operational Environment of DD-963
Class Ships; Fortenbaugh, R. L.; September 1978
AGARD Rpt 122
The Influence of Drag Characteristics on the Choice of Landing Approach
Speeds; Lean, D. and R. Eaton; 1957
AGARD Rpt 357
Some Low-Speed Problems of High-Speed Aircraft; Spence, A. and D.
Lean, 1961
AGARD Rpt 372
Theory of the Flight of Airplanes in Isotropic Turbulence - Review and
Extension; Etkin, B.; April 1961
AGARD Rpt 420
Flight Measurements of the Influence of Speed Stability on the Landing
Approach; Staples, K. J.; 1963
AGARD-AR-82
The Effects of Buffeting and Other Transonic Phenomena on Maneuvering
Combat Aircraft; Hamilton, B. I. L.; July 1975
AGARD-AR- 134
Technical Evaluation Report on the Flight Mechanics Panel Symposium on
Stability and Control; Chalk, C. R.; January 1979
AGARD-CP-17
AGARD Stability and Control Meeting, September 1966
“Flying Qualities Criteria Problems and Some Proposed Solutions,"
Carlson, John W. and Richard K. Wilson
“Pilot-Induced Instability;" A'Harrah, R. C. and R. F. Siewert
AGARD-CP- 119
Stability and Control; "Flight Simulation - A Significant Aid In Aircraft
Design;' A'Harrah, R. C.; April 1972
AGARD-CP-199
Stall/Spin Problems in Military Aircraft; June 1976
AGARD-CP-235
Dynamic Stability Parameters; “Aircraft Stability Characteristics at High
Angle of Attack;" Kalviste, J.; November 1978
AGARD-CP-249
Piloted Aircraft Environment Simulation Techniques, "Handling Qualities of
a Simulated STOL Aircraft in Natural and Computer-Generated Turbulence
and Shear;" Sinclair, S. R. M. and T. C. West; October 1978
AGARD-CP-260
Proceedings of AGARD Flight Mechanics Panel Symposium on Stability
and Control, September 1978
"Are Today's Specifications Appropriate for Tomorrow's Airplanes?"
A'Harrah, R. C., J. Hodgkinson, and W. J. LaManna
“Flying Qualities and the Fly-by-Wire Aeroplane;" Gibson, J. C.
"L-1011 Active Controls Design Philosophy and Experience;"
Urie, David M.
AGARD-CP-319
Combat Aircraft Maneuverability; "The Military Flying Qualities
Specification, a Help or a Hindrance to Good Fighter Design?" A'Harrah,
Ralph C. and Robert J. Woodcock; December 1981
Supersedes pages 62-75 of
MIL-STD-1797A
74
MIL-STD-1797A
APPENDIX A
AGARD-CP-333
Criteria for Handling Qualities in Military Aircraft; "Simulation for Predicting
Flying Qualities;" Reynolds, P. A.; June 1982
AGARD-CP-452
Flight Test Techniques; “GRATE – A New Flight Test Tool for Flying
Qualities Evaluation,” Koehler, et al; DFVLR; Jul 1989
AGARD-CP-519
Recent Advances in Long Range and Long Endurance Operation of
Aircraft; “Aerial Refueling Interoperability from a Receiver Flying Qualities
Perspective”
NATO Rpt 408A
Recommendations for V/STOL Handling Qualities; October 1964
NRC of Canada
LTR-FR- 12
Rpt
A Flight Investigation of Lateral-Directional Handling Qualities of V/STOL
Aircraft in Low Speed Maneuvering Flight; Doetsch, K. H., et al ; 15 August
1969
RAE Aero. 2504
Problems of Longitudinal Stability Below Minimum Drag, Speed, and
Theory of Stability Under Constraint; Neumark, S.; 1953
RAE Aero. 2688
A Review of Recent Handling Qualities Research, and Its Application to the
Handling Problems of Large Aircraft. Part I: Observations on Handling
Problems and Their Study. Part II: Lateral-Directional Handling; Bisgood, P.
L.; June 1964
RAE TM-FS-46
Developments in the Simulation of Atmospheric Turbulence; Tomlinson, B.
N.; September 1975
RAE TR-68140
Control Characteristics of Aircraft Employing Direct Lift Control, Pinsker, W.
J. G.; May 1968
RAE TR-71021
Glide Path Stability of an Aircraft Under Speed Constraint, Pinsker, \\'. J G.;
February 1971
TSS Standard 5
Supersonic Transport Aeroplane Flying Qualities; 22 May 1964
Ad Hoc Committee Report on B-58 Controllability in Flight, Wright Air Development Division, WrightPatterson AFB, OH, 2 April - 10 May, 1960
Anderson, Ronald O., A Second Analysis of B-58 Flight Control System Reliability, Flight Control
Laboratory, Wright-Patterson AFB, OH, 6 November 1962
Ashkenas, I. L. and T. Durand, "Simulator and Analytical Studies of Fundamental Longitudinal Control
Problems in Carrier Approach,” presented at AIAA Simulation for Aerospace Flight Conference, August.
1963
Behel, I. M. and W. B. McNamara, “F/A-18A High Angle of Attack/Spin Testing," 25th International Report
to the Aerospace Profession, Society of Experimental Test Pilots, September, 1981
Bureau of Naval Weapons Failure Rate Data Handbook, prepared by U. S. Naval Ordnance Laboratory;
Corona, CA (updated periodically)
Caravello, Christopher, Randal G. Joslin, Giuseppe Fristachi, Charles R. Bisbee, Steven S.
Weatherspoon, and Steven G. Henrich, Limited Flight Evaluation as a Function of Aircraft Longitudinal
Dynamics, Air Force Test Pilot School, Class 79A Final Report, December, 1979
Curry, R. E. and A. G. Sim, Unique Flight Characteristics of the AD-1 Oblique-Wing Research Airplane, J
Aircraft, v. 20, nr. 6, June, 1983
"Development of the F/A-18 Handling Qualities Using Digital Flight Control Technology," Society of
Experimental Test Pilots 1982 Report to the Aerospace Profession, 26th Annual Proceedings,
September, 1982
Supersedes pages 62-75 of
MIL-STD-1797A
75
MIL-STD-1797A
APPENDIX A
Dryden, Hugh L., “A Review of the Statistical Theory of Turbulence,” Turbulence - Classic Papers on
Statistical Theory, New York: Interscience Publishers, Inc., 1961
Etkin, B., "A Theory of the Response of Airplanes to Random Atmospheric Turbulence," J. Aero/Space
Sciences, July, 1959, 409-420
Etkin, Bernand, Dynamics of Atmospheric Flight, New York: Wiley, 1972
Etkin, Bernard, Dynamics of Flight, New York: Wiley, 1959
Finberg, Floyd, Report of the T-38 Flight Control System PIO Review Board. USAF ASD, February, 1963
Hirsch, Darrell, "Investigation and Elimination of PIO Tendencies in the Northrop T-38A," SAE Paper,
Amb, New York, July, 1964 NW Hodgkinson, J., "Equivalent Systems Approach for Flying Qualities
Specification," presented at SAE Aerospace Control and Guidance Systems Committee Meeting, Denver,
CO, 7-9 March, 1979
Hodgkinson, J., R. L. Berger, and R. L. Bear, “Analysis of High Order Aircraft/Flight Control System
Dynamics Using an Equivalent System Approach," presented at 7th Annual Pittsburgh Conference on
Modeling and Simulation, 26-27 April, 1976
Hodgkinson, J., W. J. LaManna, and J. L. Heyde, "Handling Qualities of Aircraft with Stability and Control
Augmentation Systems - A Fundamental Approach," J. R. Ae. S., February, 1976
Houbolt, John C., "Atmospheric Turbulence," AIAA J., Vol. II, No. 4, April, 1973, 421-437
"Industry Observer,” Aviation Week and Space Technology, 1 April, 1968, 13
Jacobson, Ira D. and Dinesh S. Joshi, “Investigation of the Influence of Simulated Turbulence on
Handling Qualities," J. Aircraft, Vol.14, No. 3, March 1977, 272-275
Jones, J. G., "Modelling of Gusts and Wind Shear for Aircraft Assessment and Certification,” Royal
Aircraft Establishment, Paper prepared for CAARC Symposium on Operational Problems, India, October,
1976
Lappe, V. Oscar and Ben Davidson, “On the Range of Validity of Taylor's Hypothesis and the Kilmogoroff
Spectral Law," J. Atmos. Sciences, Vol. 20, November, 1963
Lappe, V. Oscar, “Low-Altitude Turbulence Model for Estimating Gust Loads on Aircraft," J Aircraft, Vol. 3,
No. 1, Jan - Feb, 1966
Lumley, John L. and Hans A. Panofsky, The Structure of Atmospheric Turbulence, New York:
Interscience Publishers, Inc., 1964
McRuer, Duane, Irving Ashkenas, and Dunstan Graham, Aircraft Dynamics and Automatic Control,
Princeton University Press, 1973
Mitchell, David G. and Roger H. Hoh, “Low-Order Approaches to High-Order Systems: Problems and
Promises," J. Guidance, Control, and Dynamics, Vol. 5, No. 5, Sept - Oct 1982, 482-489
Morgan, Clifford T., Jesse S. Cook, Alphonse Chapanis, and Max W. Lund, eds., Human Engineering
Guide to Equipment Design, New York: McGraw-Hill, 1963
Morgan, Len, “Out for a Spin," Flying, February, 1982
Neal, T. Peter, "Influence of Bobweights on Pilot-Induced Oscillations," J. Aircraft, September, 1971
Otnes, R. K. and L. Enochson; Applied Time Series Analysis, Vol. 1. Basic Techniques; New York- WileyInterscience; 1978
Supersedes pages 62-75 of
MIL-STD-1797A
75a
MIL-STD-1797A
APPENDIX A
Perkins, Courtland D. and Robert E. Hage, Airplane Performance Stability and Control, New York- Wiley
1949
"Proposals for Revising MIL-F-8785B, 'Flying Qualities of Piloted Airplanes'," AFFDL-FGC Working,
Paper, February, 1978
Rediess, H. A., D. L. Mallick, and D. T. Berry, Recent Flight Test Results on Minimum Longitudinal
Handling Qualities for Transport Aircraft, presented at the FAUST VIII Meeting, Washington, D.C.,
January 1981
Richards, D. and C. D. Pilcher, "F/A-18A Initial Sea Trials," SETP Cockpit, April/May/June, 1980
Sammonds, R. I., W. E. McNeill, and J. W. Bunnell, "Criteria for Side-Force Control in Air-to-Ground
Target Acquisition and Tracking," J. Aircraft, v. 19, nr. 9, September, 1982
Scott, W. B., “Reengined KC-135 Shows Performance Gains in Test," Aviation Week & Space
Technology, v. 118, nr. 8, McGraw-Hill, February 21, 1983
Stengel, R. F. and G. E. Miller, “Pilot Opinions of Sampling Effects in Lateral-Directional Control,"
presented at 16th Annual Conference on Manual Control, Cambridge, MA, May, 1980
Tentative Airworthiness Objectives and Standards for Supersonic Transport Design Proposals, Flight
Standards Service, FAA, 15 August, 1963
Van Patten, Robert E., Investigation fo the Effects of gy. and gz on AFTI/F-16 Control Inputs. Restraints
and Tracking Performance, Interim USAF AMRL Technical Report, August, 1981
von Karman, Theodore, “Progress in the Statistical Theory of Turbulence," Turbulence - Classic Papers
on Statistical Theory, New York: Interscience Publishers, Inc., 1961
Supersedes pages 62-75 of
MIL-STD-1797A
75b
MIL-STD-1797A
APPENDIX A
30. DEFINITIONS
3.1 Aircraft classification and operational missions. For the purpose of this standard, the
aircraft specified in this requirement is to accomplish the following missions:____________. The
aircraft thus specified will be a Class aircraft. The letter -L following a class designation
identifies an aircraft as land-based, carrier-based aircraft are similarly identified by -C. When no
such differentiation is made in a requirement, the requirement applies to both land-based and
carrier-based aircraft.
REQUIREMENT RATIONALE (3.1)
The very reason for procuring aircraft is to perform one or more missions. The class designation
is used in the handbook to help particularize the requirements according to broad categories of
intended use.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 1.3, 1.3.1 and 3.1.1.
Missions
The standard needs a specific mission statement to furnish guidance for interpreting qualitative
requirements as well as for consistent selection of quantitative requirements. Unfortunately, the
word "mission" is used in several contexts not only in this standard, but throughout the writings
pertinent to acquiring a new weapon system. In the broadest sense, "operational missions"
applies to classifying the aircraft as fighter, bomber, reconnaissance, etc., or to "accomplishing
the mission" of bombing, strafing, etc. In 3.1 the object is to introduce to the designer in general
terms the function of the vehicle he is to design. It should be sufficient for the procuring activity
to refer to those paragraphs of the System Specification and Air Vehicle Specification to define
the overall performance requirements, the operational requirements, employment and
deployment requirements.
The operational missions considered should not be based on just the design mission profiles.
However, such profiles serve as a starting point for determining variations that might normally
be expected in service, encompassing ranges of useful load, flight time, combat speed and
altitude, in-flight refueling, etc., to define the entire spectrum of intended operational use.
"Operational missions" include training and ferry missions.
The intended use of an aircraft must be known before the required configurations, loadings, and
the Operational Flight Envelopes can be defined and the design of the aircraft to meet the
requirements of this standard undertaken. It additional missions are foreseen at the time the
detail specification is prepared, ft is the responsibility of the procuring activity to define the
operational requirements to include these missions. Examples of missions or capabilities that
have been added later are in-flight refueling (tanker or receiver), aerial pickup and delivery, lowaltitude penetration and weapon delivery, and ground attack for an air-superiority fighter or vice
versa.
Once the intended uses or operational missions are defined, a Flight Phase analysis of each
mission must be conducted. With the Flight Phases established, the configurations and loading
states which will exist during each Phase can be defined. After the configuration and loading
states have been defined for a given Flight Phase, Service and Permissible Flight Envelopes
can be determined and Operational Flight Envelopes more fully defined.
.
REPRINTED WITHOUT CHANGE
76
MIL-STD-1797A
APPENDIX A
4.1.6 Aircraft Normal States.
The contractor shall define and tabulate all pertinent items to describe the Aircraft Normal
States (no component or system failure) associated with each of the applicable Flight Phases.
This tabulation shall be in the format and use the nomenclature of table 11. Certain items, such
as weight, moments of inertia, center-of-gravity position, wing sweep, or thrust setting may vary
continuously over a range of values during a Flight Phase. The contractor shall replace this
continuous variation by a limited number of values of the parameter in question which will be
treated as specific States, and which include the most critical values and the extremes
encountered during the Flight Phase in question.
REQUIREMENT RATIONALE (4.1.6)
Definition of normal aircraft states is basic to application of the flying qualities requirements.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.1.6.1 and 4.2.
It is possible that hems not normally considered, such as setting or automatic operation of engine bypass
doors, can affect flying qualities.
The contractor is required to define the Aircraft Normal States for each applicable Flight Phase, in the
format of table ll. If the position of any particular design feature can affect flying qualities independently of
the items in table 11, its position should be tabulated as well. Initially, variable parameters should be
presented in discrete steps small enough to allow accurate interpolation to find the most critical values or
combinations for each requirement; then those critical cases should be added. As discussed under 4.1.1
through 4.1.3, center-of-gravity positions that can be attained only with prohibited, failed, or
malfunctioning fuel sequencing need not be considered for Aircraft Normal States.
REQUIREMENT LESSONS LEARNED
5.1.6 Aircraft Normal States-verification.
The contractor shall furnish a list of Aircraft Normal States in accordance with the Contract Data
Requirements List.
VERIFICATION RATIONALE (5.1.6)
Definition of normal aircraft states is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
Definition of normal aircraft states is basic to application of the flying qualities requirements.
VERIFICATION LESSONS LEARNED
REPRINTED WITHOUT CHANGE
107
MIL-STD-1797A
APPENDIX A
4.1.6.1 Allowable Levels for Aircraft Normal States.
Flying qualities for Aircraft Normal States within the Operational Flight Envelope shall be Level 1. Flying
qualities for Aircraft Normal States within the Service Flight Envelope but outside the Operational Flight
Envelope shall be Level 2 or better. To account for the natural degradation of pilot-vehicle performance
and workload in intense atmospheric disturbances, the requirements of 4.1.6.1 through 4.1.6.3 are
adjusted according to 4.9.1.
4.1.6.2 Flight outside the Service Flight Envelopes.
From all points in the Permissible Flight Envelopes and outside the Service Flight Envelopes, 4 shall be
possible readily and safety to return to the Service Flight Envelope without exceptional pilot skill or
technique. The requirements on flight at high angle of attack, dive characteristics, dive recovery devices
and dangerous flight conditions shall also apply in all pertinent parts of the Permissible Flight Envelopes.
4.1.6.3 Ground operation.
Some requirements pertaining to taxiing, takeoffs, and landing involve operation outside the Operational,
Service, and Permissible Flight Envelopes, as at VS or on the ground. When requirements are stated at
conditions such as these, the Levels shall be applied as it the conditions were in the Operational Flight
Envelopes.
REQUIREMENT RATIONALE (4.1.6.1 - 4.1.6.3)
Levels of flying qualities as indicated in 3.3 apply generally within the Operational and Service
Flight Envelopes. Some basic requirements, generally qualitative in nature, apply in both the
Operational and Service Flight Envelopes. Provision must also be made for expected and
allowable operation outside these envelopes.
REQUIREMENT GUIDANCE
The related MIL-F-4785C requirements are paragraphs 3.1.10, 3.1.10.1, 3.1.10.3.1, 3.1.10.3.2,
3.1.10.3.3, 3.8.3, and 3.8.3.1.
Aircraft Normal States include both all-up operation and degradations/failures that are sufficiently
probable to be considered Normal. See 4.1.7 and 4.1.7.1 for guidance on the latter.
Note that flying qualities which "warrant improvement" according to figure 6 nevertheless meet all the
requirements if they only occur outside the Operational Flight Envelope.
Where Levels are not specified, care should be taken in selecting requirements from this handbook that
will not overburden the designer. We have tried to keep the impact of 4.1.6.1 in mind in writing the
recommended material to fill in the blanks, but qualitative words such as "objectionable" must be taken in
the context of relevance to operational use.
Since there are few requirements in Aircraft Failure States outside the Service Flight Envelope, implicit
assumptions for 4.1.6.2 are that:
Failures at these conditions are very rare, or
Not-so-rare failures at these conditions are manageable
Given one or more failures within the Service Flight Envelope which would have serious
consequences beyond, at a minimum the crew would be warned away from danger
(4.1.8).
Similar assumptions apply for 4.1.6.3. In any given case, their validity will need to checked.
REQUIREMENT LESSONS LEARNED
5.1.6.1 Allowable Levels for Aircraft Normal States - verification.
Verification shall be by analysis, simulation, and test. Final verification shall be by demonstration in the
performance of the following tasks:
Supersedes page 108 of MIL-STD-1797A
108
MIL-STD-1797A
APPENDIX A
5.1.6.2 Flight outside the Service Flight Envelopes - verification.
Verification shall be by analysis, simulation and test. Final verification shall be by demonstration in the
performance of the following tasks:
5.1.6.3 Ground operation - verification.
Verification shall be by analysis, simulation and test. Final verification shall be by demonstration in the
performance of the following tasks:
VERIFICATION RATIONALE (5.1.6.1 - 5.1.6.3)
Compliance with quantitative, open-loop (pilot-out-of-the-loop) requirements can be shown by analysis
with data derived from flight test. However, compliance with only the quantitative, open-loop requirements
does not guarantee that the required Levels of flying qualities have been achieved. The quantitative,
open-loop requirements are based on historical data and research experiments with ground-based and
in-flight simulators. Obviously, this data base has not evaluated each and every one of the infinite
possible combinations of aircraft physical and dynamic characteristics at every possible flight condition in
every possible task. The criteria in the quantitative, open-loop requirements are based on interpolation
between and extrapolation beyond the configurations, flight conditions, and tasks that have been
evaluated in the existing data base. Given the degree of generalization in this standard (all possible
aircraft configurations generalized into four Classes, all possible tasks generalized into three Flight Phase
Categories, and all possible flight conditions generalized into three Flight Envelopes) a certain amount of
discrepancy between Level boundaries and pilot-observed flying qualities is to be expected occasionally.
Furthermore, while the most significant factors affecting flying qualities have been identified and Level
boundaries established for them, significant factor interactions, which have not been as well identified to
date, may also cause discrepancies between Level boundaries and observed flying qualities. Other
sources of such discrepancies are pilot and task variabilities. The pilots used to evaluate a configuration
in a flight test program or a research experiment, and the levels of task performance required for desired
and adequate ratings, will certainly skew the results. Therefore, in order to insure that the aircraft has
achieved the required Levels of flying qualities, the aircraft must be evaluated by pilots in high-gain,
closed-loop tasks. (In the context of this document, high-gain task means a wide-bandwidth task, and
closed-loop means pilot-in-the-loop.) For the most part, these tasks must be performed in actual flight.
However, for conditions which are considered too dangerous to attempt in actual flight (i.e., certain flight
conditions outside of the Service Flight Envelope, flight in severe atmospheric disturbances, flight with
certain Failure States and Special Failure States, etc.), the closed-loop evaluation tasks can be
performed on a simulator.
During the requirements definition process, the procuring agency, together with the contractor(s) and the
responsible test organization, should select several closed-loop tasks with which to evaluate the aircraft in
flight test. During system development, ground-based and in-flight simulations should be used to get an
initial appraisal of how well the aircraft will perform the tasks in flight. The simulations can also be used to
train the test pilots and refine the tasks, performance objectives, and test procedures. Handling qualities
evaluation during flight test should consist of four parts: 1) "open-loop" tasks such as steps, doublets, and
frequency sweeps for parameter identification to compare aircraft dynamic response to the open-loop
requirements, 2) capture tasks to familiarize the pilot with aircraft response and capture characteristics, 3)
Handling Qualities During Tracking (HQDT) for initial closed-loop handling qualities evaluation, and 4)
"operational” closed-loop tasks to obtain Cooper-Harper ratings. The distinction between HQDT and
"operational” tasks is discussed in 5.1.11.6 Verification Guidance.
VERIFICATION GUIDANCE
A wide variety of closed-loop tasks have been developed for the evaluation of aircraft flying qualities.
Recommended tasks for the evaluation of flying qualities for Flight Phase Category A include air-to-air
and air-to-ground tracking, particularly the well-defined set of tasks known as HQDT, aerial refueling,
close formation flying, and captures. Recommended tasks for the evaluation of flying qualities for Flight
Phase Category C are tracking tasks, including HQDT, close formation flying, precision landings (with and
without vertical and lateral offsets), takeoffs, and captures. More detailed discussions of each of these
recommended tasks will be found in subsequent paragraphs of this section and 5.1.11.6 Verification
Guidance. Other tasks which are critical to the mission that the aircraft is intended for may also be used
as evaluation tasks, such as terrain-following, weapons delivery, or LAPES.
NEW PAGE
108a
MIL-STD-1797A
APPENDIX A
There are no recommended tasks for Flight Phase Category 8 because this Flight Phase Category
generally consists of low-gain (low-bandwidth) tasks. Possible flying qualities problem in this Flight Phase
Category will normally be exposed in the more demanding tasks for Flight Phase Categories A and C,
and by normal operations during the flight test program. Thus, special tasks for this Flight Phase
Category are not normally considered necessary. However, Flight Phase Category B tasks are normally
of much longer duration than the tasks in the other Categories. Pilot fatigue may become a significant
factor in certain mission critical Category B tasks, in which case an evaluation of this kind of task might be
required.
Proof of compliance in these demonstration tasks will consist of pilot comments and Cooper-Harper (C-H)
ratings. For Level 1, pilot comments must indicate satisfaction with aircraft flying qualities, with no worse
than "mildly unpleasant" deficiencies, and median C-H ratings must be no worse than 3.5 in calm air or in
light atmospheric disturbances. For Level 2, pilot comments must indicate that whatever deficiencies may
exist, aircraft flying qualities are still acceptable, and median C-H ratings must be no worse than 6.5 in
calm air or light atmospheric disturbances. For Level 3, pilot comments must indicate that the aircraft is at
least controllable despite any flying qualities deficiencies, and median C-H ratings must be no worse than
9.5 in calm air or light atmospheric disturbances. In moderate to severe atmospheric disturbances pilot
comments and C-H ratings must comply with the requirements of 4.9.1.
Actual task performance is not recommended for use as proof of compliance because it is even more
subject to pilot variability than pilot comments and C-H ratings. The performance objectives suggested in
the tasks described below are not intended for use as proof of compliance, but, rather, for use with the CH scale. Specific definitions of desired and adequate performance objectives reduce pilot variability by
insuring that all of the evaluation pilots attempt to achieve the same level of performance. In the
performance objectives suggested below, adequate performance is set at a level sufficient to successfully
perform similar tasks in operational service. Desired performance is set at a more demanding level to
insure that system deficiencies are exposed. Although task performance is not recommended as proof of
compliance, task performance should be recorded and analyzed by the flight test engineers to insure that
pilot ratings are reasonably consistent with the level of performance achieved and that all pilots seem to
be attempting to achieve the same level of performance.
The evaluation of aircraft flying qualities is basically a subjective science, and human variability makes
analysis of the results a difficult proposition. Nevertheless, there are steps that can be taken to reduce
variability in the results and insure a good evaluation. First of all, it is absolutely necessary that more than
one evaluation be made for each test condition. Studies of inter-pilot C-H rating variability have indicated
that three pilots is the minimum number of pilots for an adequate evaluation (CAL Report No. TB-1444-F1 and NADC-85130-60). More pilots will increase confidence in the results, but NADC-85130-460 further
demonstrated that the point of diminishing returns was reached at about six. Therefore, the
recommended number of pilots per test condition is three to six. Careful selection of the evaluation pilots
will also reduce the variability in results. All of the evaluation pilots must be test pilots trained in the use of
the C-H scale and they all must be experienced in the Class of aircraft under evaluation. Furthermore, for
acquisition purposes, it is highly desirable that at least half of the evaluation pilots be military-employed
test pilots. (Use of operational pilots to evaluate the aircraft during the development effort can often
provide additional insights into the handling qualities. Such evaluations are strongly encouraged.
However. for the purpose of demonstrating compliance with this requirement, the evaluation pilots should
be trained test pilots.)
In order to insure that all of the pilots attempt to achieve the same level of performance, and thus insure
consistency and reduce the effects of pilot variability, it is extremely important to explicitly define the
desired and adequate levels of task performance to be used for the C-H ratings. Best results are
achieved with task performance defined in terms of quantifiable objectives which the pilot can readily
observe himself in real time. Furthermore, consideration must be given to defining objectives that can
also be recorded on some medium so that the flight test engineer can confirm that pilot ratings are
reasonably consistent with task performance. Defining quantifiable and recordable task performance
objectives and setting appropriate levels of desired and adequate performance is the most difficult part of
planning the flying qualities evaluation. Guidance on task objectives for each of the recommended tasks
is given in subsequent paragraphs in this section and lessons learned from past experience are provided
in Verification Lessons Learned.
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Another method to reduce the effects of variability is to use the “long-look” evaluation technique. In this
technique the pilot continues or repeats the task until he is confident of his evaluation before he assigns a
C-H rating. (As opposed to doing it once and assigning a rating.) The "long-look" approach allows the pilot
a more extensive appraisal of the test condition. it allows him to weed out the effects of unique events in a
single run. It allows him to get over the learning curve, and it allows him to clear his memory of
characteristics he may have observed in evaluation of a preceding condition. Although the C-H rating is
given only after multiple runs, pilot comments must be provided during and after each run. In order to
insure that variability is not introduced by pilots doing different numbers of repeat runs, the recommended
procedure is to specify a minimum number of runs to be performed before a rating can be given, and
allow the pilot to make additional runs it he feels it is necessary.
Pilot comments should be considered the most important data. A C-H rating is only a summary of
observed flying qualities characteristics into a single number. Pilot comments identity the specific
deficiencies, if any, that must be corrected. Furthermore, the "long-look" technique filters the effects of
deficiencies on the C-H rating because, over several runs, the pilot learns to compensate for some
deficiencies. Since pilot comments are given for every run, the comments will identity all observed
deficiencies, even those which can be compensated for. Comments on succeeding runs provide guidance
on the pilot's ability to compensate for the deficiencies and the final C-H rating indicates the relative
significance of these deficiencies. Therefore, pilot comments must be recorded and analyzed for every
test run.
Time and cost constraints prohibit piloted evaluation of every task in every possible aircraft configuration
at every possible point in the flight envelope. The conditions that must be evaluated are the most
common operating conditions, operating conditions critical to the mission of the aircraft, and the worst
case conditions, especially those where the quantitative, open-loop flying qualities requirements are
violated by wide margins. For aircraft with multiple flight control modes, all mode transitions should be
evaluated at common, mission critical, and worst case conditions, especially mode switches which are
done automatically, as opposed to those deliberately switched by the pilot. Furthermore, the degradation
due to atmospheric disturbances should be demonstrated by evaluation at different levels of
disturbances. Evaluation of the effect of severe atmospheric disturbances may be performed in ground
simulation. When using simulation to predict the degradation of flying qualities due to severe atmospheric
disturbances, it will be necessary to correlate C-H ratings gathered from simulation sessions in light to
moderate turbulence with C-H ratings obtained from flight test in light to moderate turbulence for the
same task.
The following paragraphs discuss some recommended tasks and some suggested performance
objectives for each task. BEAR IN MIND THAT THE PERFORMANCE OBJECTIVES ARE NOT
REQUIRED AS PROOF OF COMPLIANCE. THEY ARE INTENDED FOR USE WITH THE COOPERHARPER SCALE TO REDUCE PILOT VARIABILITY. Most of these tasks are equally suitable for both the
"operational" technique and the HQDT technique (see 5.1.11.6 Verification Guidance). The following list
of tasks is not exclusive. Any closed-loop task, performed aggressively, may be used to evaluate an
aircraft's handling qualities and PIO characteristics. When developing a specification for a particular
program, the procuring agency should discuss possible tasks and performance objectives with other
procuring agencies, Wright Laboratory, AFFTC, and potential contractors.
Capture Tasks
Capture tasks are intended to evaluate handling qualities for gross acquisition as opposed to continuous
tracking. A wide variety of captures can be done provided that the necessary cues are available to the
pilot. Pitch attitude, bank angle, heading, flight path angle. Angle-of-attack, and g captures have all been
done in previous programs to evaluate different aspects of the aircraft response. These capture tasks are
done almost precognitively by the pilot and are usually over so quickly that they do not lend themselves
well to use with the Cooper-Harper scale. It can be done, of course, but it is not necessary. Qualitative
comments are sufficient proof of compliance for these tasks.
These capture tasks can give the pilot a general impression of the handling qualifies of the
aircraft, but because they do not involve closed-loop tracking they do not expose all of the
problems that arise in continuous control tasks. Capture tasks should riot be used as the only
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evaluation tasks. As a minimum an off set precision landing task and some form of tracking task
should also be used. Capture tasks are
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ideal as a pre-test before performing high-gain, closed-loop tasks because they serve to
familiarize the pilot With the aircraft response before attempting the more difficult (and
sometimes more dangerous) high-gain, closed-loop tasks. If hazardous motions result from the
capture tasks at any flight condition, closed-loop tasks should not be attempted at that flight
condition, and the aircraft should be considered to have failed this requirement at that condition.
For pitch captures, the aircraft is trimmed about a specified flight condition. The pilot aggressively
captures 5° pitch attitude (or 10° if the aircraft is already trimmed above 5°). He then makes a series of
aggressive pitch captures of 5° increments in both directions. He then continues this procedure with 10°
increments in both directions, and then with 15° increments in both directions. Aircraft with more
capability can continue the procedure with larger pitch excursions. It possible, the initial conditions for
each maneuver should be such that the aircraft will remain within ±1000 feet and ±10 knots of the
specified flight condition during the maneuver, however, large angle captures at high-speed conditions
will inevitably produce larger speed and altitude changes. If the aircraft should get too far from the
specified condition during a task, it should be retrimmed about the specified condition before starting the
next maneuver.
The other kinds of captures are usually done in a similar manner, with some minor differences. G
captures are usually done from a constant-g turn and the increments are usually ±1 g, ±2g, and ±3g. and
larger increments if the aircraft has greater capability. Heading captures ran be used to evaluate the yaw
controller alone (usually small heading changes of 5° or less) or to evaluate coordination of yaw and roll
controllers (larger heading changes).
Bank angle captures are also commonly done using bank-to-bank rolls. Starting from a 150 bank angle,
the pilot aggressively rolls and captures the opposite 15° bank angle (total bank angle change of 30°). He
then rolls back and captures 150 bank in the original direction. This procedure should continue for a few
cycles. The procedure is then repeated using 30° bank angles, and then repeated again using 45° bank
angles. Aircraft with more capability can continue the procedure with larger bank angles. A variation of
this is to capture wings-level from the initial banked condition. Four-point and eight-point rolls, standard
aerobatics maneuvers, are also good tests of roll control.
Air-to-Air Tracking
The air-to-air tracking task consists of two phases: gross acquisition and fine tracking. Gross acquisition
evaluates the ability to point and capture with mode rate-amplitude inputs. Fine tracking evaluates
continuous closed-loop controllability. Two different kinds of targets have been used successfully in
handling qualities evaluations: a real target aircraft and a target generated by a HUD (Head-Up Display).
If a real target is used, there are several possible target maneuvers which have been used in handling
qualities evaluations in the past. In all cases the target aircraft begins the maneuver from straight and
level flight in front of the evaluation aircraft at a specific flight condition. Throughout the maneuver the
evaluation aircraft should remain within ±1000 feet of the test altitude and within ±50 knots of the test
airspeed.
The maneuver most commonly used is an S-turn. The target aircraft initiates a level turn at a specified
load factor. After a specified time period the target unloads, reverses, and begins a level turn in the
opposite direction at the specified load factor. After a specified time period, the maneuver is terminated.
For gross acquisition, the evaluation pilot should allow the target aircraft to achieve a certain
amount of angular displacement before he initiates his maneuver to acquire the target. Some
programs have stated the angular displacement explicitly (for example, 100 mils or 30°). Other
programs have used the point at which the target crosses the canopy bow to initiate gross
acquisition. Yet another option is to allow a specified amount of time between the target aircraft
turn and the evaluation aircraft turn (3-4 seconds). Commonly used performance objectives for
gross acquisition are time to acquire, the number of overshoots, and the size of the overshoots.
Acquisition is defined as bringing the pipper (or whatever the pilot is using to track with) within a
certain radius of some specified point on the target (tail pipe, fuselage/wing junction, canopy,
etc.). Time to acquire is the time it takes to bring the pipper within this radius and keep it there.
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APPENDIX A
Time to acquire is a difficult objective to recommend specific values for in a general standard
because it is not only a function of handling qualities, but also a function of handling
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APPENDIX A
of the size of the initial angular displacement between pipper and target point and of the maximum pitch
rate performance of the aircraft under evaluation. Programs must consider both of these when
determining what time to acquire to specify. An overshoot is when the pipper moves past the target point
by some amount of angular displacement. Some suggested performance objectives are given in table
LVIll.
An illustration of overshoot is shown in the sketch below, which shows the time histories of three
theoretical gross acquisitions. Assume the desired criteria are: no more than one overshoot greater than
5 mils and no overshoots greater than 10 mils. The thick continuous line would fail this criteria because it
has one overshoot greater than 10 mils (at about 0.5 seconds). The thin continuous line also fails this
criteria because it has two overshoots greater than 5 mils (around 0. 5 seconds and 0.9 seconds). The
dashed line meets this criteria because it has only one overshoot that exceeds 5 mils but that overshoot
is less than 10 mils.
For fine tracking, the time between reversals should allow for time to acquire plus time for extended
tracking. A minimum of at least 20 seconds between reversals is recommended. The nominal range
between the target aircraft and the evaluation aircraft should be about 1500 feet, with excursions of no
more than ±500 feet from nominal. The performance objective for fine tracking is to keep the pipper within
a certain radius of the target point for a large percentage of the tracking time. Some suggested
performance objectives are given in table LVIII.
For the long-look technique, this maneuver should be repeated a few times before giving a C-H rating. On
a ground-based simulator the sequence of turns can continue uninterrupted until the evaluation pilot is
confident that he has a good evaluation of the aircraft. This evaluation should be conducted at different
airspeeds, different altitudes, and with different load factors throughout the Operational and Service Flight
Envelopes.
Another common target maneuver is the wind-up turn. In the wind-up turn, the target aircraft begins a turn
and slowly and smoothly increases the load factor to a specified maximum load factor. The target aircraft
should attempt to maintain a specified rate of g increase, about .2 g/sec is recommended. The maneuver
is terminated shortly after the maximum load factor is reached. Gross acquisition in this maneuver is
similar to that for the S-turns. For fine tracking, the rate of g increase should allow sufficient time for gross
acquisition and extended tracking time. For the long-look technique this maneuver should be repeated a
few times. This evaluation maneuver should be initiated from various altitudes and airspeeds throughout
the Operational and Service Flight Envelopes.
Other target-tracking maneuvers that have been used in the past are discussed in Verification Lessons
Learned.
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APPENDIX A
TABLE LVIII. Suggested performance objectives for various evaluation tasks.
Suggested Tasks
Suggested Performance Objectives
Air-to-Air and
Desired Performance
Air-to-Ground Tracking:
Time to acquire: TBD
Gross Acquisition
Overshoots: no more than one greater than 5 mils. none to exceed 10 mils
No PIO
Adequate Performance
Time to acquire: TBD
Overshoots: no more than two greater than 5 mils, none to exceed 20 mils
Air-to-Air and
Desired Performance
Air-to-Ground Tracking:
Keep the pipper within 5 mils of the target point for three continuous
seconds
Fine Tracking
No PIO
Adequate Performance
Keep the pipper within 10 mils of the target point for three continuous
seconds
Close Formation
Desired Performance
Excursions no greater than ±2 feet from the formation position
No PIO
Adequate Performance
Excursions no greater than ±4 feet from the formation position
Aerial Refueling:
Desired Performance
Boom Tracking
Keep pipper within 5 mils of the boom nozzle for at least 50% of the tracking
time
No PIO
Adequate Performance
Keep pipper within 10 mils of the boom nozzle for at least 50% of the
tracking time
Aerial Refueling:
Desired Performance
Probe-and-drogue
Hook-up without touching basket webbing in at least 50% of the attempts
No PIO
Adequate Performance
Hook-up in at least 50% of attempts
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APPENDIX A
Offset Precision
Desired Performance
Landing: Approach
Flightpath control: Remain within ±1 degree of glideslope angle or ± 1 dot
2
on ILS
Airspeed control: Maximum of 5 knots above approach speed, minimum
TBD
No PIO
Adequate Performance
Flightpath control: Remain within ±2 degrees of glideslope angle or ±1 dot
on ILS
Airspeed control: Maximum of 10 knots above approach speed, minimum
TBD, but not less than Vstall
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APPENDIX A
TABLE LVIII. Suggested performance objectives for various evaluation tasks - Cont'd.
Suggested Tasks
Suggested Performance Objectives
Offset Precision
Desired Performance
Landing: Touchdown
Touchdown zone: within ±25 feet of aimpoint laterally, within -100 to +400
feet of aimpoint longitudinally
(Conventional aircraft)
Speed at touchdown: maximum of 5 knots above landing speed, minimum
TBD
Attitude at touchdown: TBD
Sink rate at touchdown: TBD
No PIO
Adequate Performance
Touchdown zone: within ±50 feet of aimpoint laterally, within -250 to +750
feet of aimpoint longitudinally
Speed at touchdown: maximum of 10 knots above landing speed, minimum
TBD
Attitude at touchdown: TBD
Sink rate at touchdown: TBD
Offset Precision
Desired Performance
Landing: Touchdown
Touchdown zone: within ±10 feet of aimpoint laterally, within -25 to +75 feet
of aimpoint longitudinally
(STOL aircraft)
Speed at touchdown: maximum of 2 knots above landing speed, minimum
TBD
Attitude at touchdown: TBD
Sink rate at touchdown: TBD
No PIO
Adequate Performance
Touchdown zone: within ±25 feet of aimpoint laterally, within -100 to +400
feet of aimpoint longitudinally
Speed at touchdown: maximum of 5 knots above landing speed, minimum
TBD
Attitude at touchdown: TBD
Sink rate at touchdown: TBD
Offset Precision
Desired Performance
Landing: Rollout and
Keep the nosewheel within ±10 feet of the runway centerline
Takeoff Roll
No PIO
Adequate Performance
Keep the nosewheel within ±25 feet of the runway centerline
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APPENDIX A
Takeoff Rotation
Desired Performance
Attitude control: Keep within ±1 degree of takeoff attitude
Overshoots: no more than one overshoot, not to exceed TBD degrees
No PIO
Adequate Performance
Attitude control: Keep within ±2 degrees of takeoff attitude
Overshoots: no more than one overshoot, not to exceed TBD degrees
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APPENDIX A
TABLE LVIII. Suggested performance objectives for various evaluation tasks - Cont'd.
Suggested Tasks
Suggested Performance Objectives
Takeoff Climbout
Desired Performance
Flightpath control: Keep within ±1 degree of specified climbout angle
Groundtrack: Keep aircraft within ±10 feet of runway centerline or within ±2
degrees of runway heading
No PIO
Adequate Performance
Flightpath control: Keep within ±2 degrees of specified climbout angle, but
not less than 0°
Groundtrack: Keep aircraft within ±25 feet of runway centerline or within ±5
degrees of runway heading
An alternative to using a real target aircraft is to do a HUD tracking task. In this task, a target symbol
(tracking bars or a line-drawing of a target) is projected on the HUD which commands pitch and roll
changes that the evaluation pilot must follow. The pitch and roll commands can be combinations of steps
and ramps, a smoothly-varying function (such as a sum-of-sines) or a simulated target aircraft maneuver
(such as those described above). The sequence of pitch and roll commands should be designed so as to
keep the aircraft within =1000 feet of the test attitude and within ±50 knots of the test airspeed. The
sequence should be long enough and complex enough that the pilot cannot learn to anticipate the
commands. Some example sequences that have been used before are discussed in Verification Lessons
Learned.
Air-to-Ground Tracking
The air-to-ground tracking task has two phases: gross acquisition and fine tracking. Gross acquisition
evaluates the ability to switch from one target to another. Fine tracking evaluates the ability to
continuously track a target.
For this task, the aircraft flies at a specified glideslope and airspeed toward a group of widely-spaced
targets on the ground. The airspeed and glideslope should be representative of the intended operational
application for the aircraft. Initial range to the targets should allow time for acquisition and tracking of
several targets. The targets should be from 60 to 180 feet apart perpendicular to the flight path and
anywhere from 90 to 360 feet apart parallel to the flight path. The pilot aggressively captures the first
target and tracks it for a specified period of time (4 seconds is recommended), and then acquires and
tracks succeeding targets. The sequence of targets to be tracked should be specified in advance. As the
aircraft approaches the targets the angular displacement between the targets will increase. Therefore, at
long range the sequence should require switching between the more widely-spaced targets (from one end
of the group to the other end, for example). As the range closes the sequence should require switching
between targets which are closer together (adjacent targets). The last target switch should require a pitch
up. A minimum recovery altitude should be specified at which the pilot must abandon the task. This
minimum attitude should consider the airspeed and dive angle of the task and should allow plenty of
margin for the pilot to pull out. Suggested performance objectives are the same as those for air-to-air
tracking. Some suggested performance objectives are given in table LVIII.
Close Formation Task
While the tracking tasks put a lot of emphasis on attitude control, close formation tasks put more
emphasis on flight path control. The task consists of holding close formation with a target aircraft as it
maneuvers. Both wing and trail formations are used. The performance objective is to maintain relative
position between the target aircraft and the aircraft under evaluation. Specific objectives are difficult to
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APPENDIX A
recommend because the pilot usually uses visual alignment of some part of his aircraft against the target
aircraft to gauge his position, and, of course, this will vary with evaluation aircraft and target aircraft.
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APPENDIX A
Some suggested displacements to use for performance objectives are given in table LVIII, however,
consideration should be given to the nature of the target maneuvers. More relaxed performance
objectives should be used for the more extreme target maneuvers. A procuring agency will have to
determine for themselves what these numbers mean in terms of visual alignment with the target aircraft in
their program.
The target maneuvers used in this task are usually the same or similar to the ones used in the air-to-air
tracking task, including the ones mentioned in Verification Lessons Learned. For the long-look technique
the task should be done a few times in the wing formation and then done a few times in trail formation
and then rated. The pilot's objectives in this task are outside visual references which are difficult to gauge
and are usually not recorded by any medium (unless a video camera is specially mounted for this task or
an observer is carried in one of the aircraft).
Aerial Refueling
There are two types of aerial refueling: probe-and-drogue and flying boom. To date, evaluation tasks
associated with both types of aerial refueling have not been formalized as much as those for tracking and
precision landing tasks. Some formalized tasks which have been used in previous programs are
described below.
For boom-type aerial refueling, the most frequently used evaluation task is some type of boom tracking
task. Two approaches have been tried. The first approach is to track the nozzle of the boom with a
waterline symbol or a pipper on the HUD or windscreen. The evaluation aircraft takes station in the precontact position about 50 feet aft and down on a 30 deg line from the tanker, which maintains steady,
level flight with the boom extended. The boom may be held stationary or moved around slowly (no more
than 1 deg/sec). The tracking time should be extensive for a good evaluation: two to four minutes is
recommended. The performance objective is to keep the pipper or waterline symbol within a certain
radius of the nozzle of the boom for a large percentage of the tracking time. Some suggested
performance objectives for this approach are given in table LVIII.
In the second approach the evaluation pilot attempts to keep the end of the boom visually aligned with
some point on the tanker aircraft. The evaluation aircraft maintains the pre-contact position within about
±10 feet. In this approach the boom is held stationary. The recommended tracking time is two to four
minutes. The performance objective in this approach is to keep the end of the boom visually aligned
within a clearly discernible area on the tanker for a specified percentage of the tracking time. It is difficult
to recommend performance objectives for this approach because they will depend on the type of tanker in
use. However, an example of a project which used this approach with a KC-135 is given in Verification
Lessons Learned.
For probe-and-drogue refueling actual hook-ups have been used as evaluation tasks. In one such
program, the performance objective was the ratio of successful hook-ups to attempted hook-ups. The
task starts from the standard pre-contact position. When cleared for contact, the evaluation pilot
establishes a 3-5 knot closure rate towards the drogue and attempts to make contact. If the drogue is
successfully engaged, the evaluation pilot stabilizes for approximately 30 seconds, and then establishes a
3-5 knot separation rate to disconnect and return to the pre-contact position. It the closure rate exceeds 5
knots, the probe tip passes the outside edge of the drogue basket, the probe tips the basket, or if a
hazardous situation develops, the hook-up attempt is aborted and the evaluation pilot returns to the precontact position before making another attempt. The performance objective is a certain percentage of
successful hook-ups out of a specific number of attempts. Six to twelve is the recommended number of
attempts. Some suggested performance objectives are given in table LVIII.
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APPENDIX A
Offset Precision Landing
The standard evaluation task for approach and landing is an offset precision landing. There are up to
three phases in this task: approach, touchdown, and, sometimes, rollout. The approach phase evaluates
the ability to control flightpath, airspeed, and attitude, including gear-down transients and large amplitude
maneuvers. The touchdown phase evaluates the ability to control flightpath, airspeed, and attitude to a
precise touchdown in the presence of ground effects and through touchdown transients. Rollout evaluates
ground handling after touchdown.
The approach phase evaluation begins about a mile out on final approach, with gear and flaps up, at the
required glideslope angle, but with a lateral offset from the runway centerline of about 150 to 300 feet,
and a vertical offset from the glideslope of about 100 to 200 feet. Soon after the task begins the pilot
lowers gear and flaps to the landing configuration. The pilot maintains precise flightpath angle and
airspeed control throughout the approach phase up to the off set correction point. Some suggested
performance objectives are given in table LVIII. Determination of performance objectives for flight path
control should consider what cues are available to the pilot. Determination of performance objectives for
airspeed control must consider the margin between the recommended approach speed and Vstall.
Lateral offsets of 200 feet or less should be corrected at 150 feet AGL. Lateral offsets of more than 200
feet should be corrected at 200 feet AGL. The pilot should make an aggressive correction to the
glideslope. The correction should be completed with the wings approximately level by 50 feet AGL to
avoid the possibility of striking the ground.
The touchdown phase begins at about 50 feet AGL. The pilot attempts to put the main wheels down
inside a designated touchdown zone at a specified landing speed, attitude, and sink rate. The landing
zone should be clearly identifiable on the runway. Performance objectives are the touchdown location.
landing speed, attitude, and sink rate at touchdown. Some suggested performance objectives for
conventional aircraft are given in table LVIll. Determination of performance objectives for airspeed control
must consider the margin between the recommended landing speed and Vstall . Similarly, performance
objectives for attitude at touchdown must consider aircraft geometry (to preclude wingtip or tail strikes,
etc.) and landing gear limitations (side force limits, etc.). Landing gear limitations must also be considered
in the determination of performance objectives for sink rate at touchdown.
Performance objectives for STOL and carrier-based aircraft should be more demanding. Suggested
performance objectives for STOL aircraft are also given in table LVIII.
The rollout phase begins after touchdown. The pilot steers from the touchdown point to the runway
centerline and thereafter stays on the runway centerline while bringing the aircraft to a stop. For STOL
aircraft a stopping distance should be specified (usually set by mission performance requirements). Some
suggested performance objectives are given in table LVIII. Landing rollout should be evaluated with and
without crosswinds using rudder and nosewheel steering or differential braking. Landing rollout might also
be evaluated under various runway conditions (dry, wet, icy, patched, etc.). Ground handling qualities
should be expected to degrade with degraded runway conditions in a similar manner to the way handling
qualities degrade with atmospheric disturbances, however, at this time there is no available guidance on
how much degradation to allow under various runway conditions. The best that can be said at this time is
that ground handling should be Level 1 for normal runway conditions, and if the aircraft cannot be kept on
the runway under certain conditions, the aircraft should be considered uncontrollable for those conditions.
As is the case in evaluations with atmospheric disturbances, dangerous runway conditions should only be
tested in ground simulation.
In flight test, bringing the aircraft to a full stop on the runway on every run is in advisable due to cost and
time constraints. Therefore, for the long-look technique, approach and touchdown should be evaluated by
doing a few touch-and-goes before giving a C-H rating. Rollout should be evaluated on the last
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landing, which is brought to a complete stop. On a ground-based simulator, the aircraft should be brought
to a stop every time. On ground-based simulators, pilots tend to be "lower gain” than they are in flight. To
counter this, light random turbulence and fairly large discrete gusts should be introduced throughout the
task. In particular, a discrete gust should be introduced after the offset correction. Degradation with
atmospheric disturbances (4.9.1) should be evaluated by increasing the turbulence and the gusts and by
adding crosswinds and wind shears.
Because this task is done in close proximity to the ground it should not be attempted if other evaluations
(analysis, ground simulation, or flight test) indicate a high probability of Level 3 handling qualities or
hazardous PIO tendencies. Therefore offset precision landing tasks should be performed on a groundbased simulator and approach handling qualities should be evaluated with other in-flight tasks (such as
HQDT) before attempting actual off set precision landings in flight test. Obviously, if this task is
considered too dangerous to attempt, the aircraft is considered to have failed this requirement.
Takeoff
Takeoff tasks have not been done as often as landing and tracking tasks, so there is little practical
experience on which to base the task recommended here. The task consists of three phases: takeoff roll,
rotation, and climb-out. The takeoff roll evaluates ground handling from brake release to takeoff rotation.
The rotation phase evaluates ability to control attitude during takeoff. The climb-out phase evaluates
ability to control flightpath after takeoff, including leaving ground effect and gear transients.
The takeoff roll begins from takeoff condition at the end of the runway. The pilot advances the throttles to
a specified setting and releases the brakes. The task is to track the runway centerline as the aircraft
accelerates. The suggested performance objectives for this phase are the same as those for landing
rollout. Some suggested performance objectives for the takeoff roll are given in table LVIII.
At a specified speed the pilot briskly rotates the aircraft to takeoff attitude. Performance objectives in this
phase are attitude control, number of overshoots, and size of overshoots. An overshoot in this case is
defined as any deviation above the specified takeoff attitude. The purpose of the overshoot limit is to
prohibit over-rotation. Some suggested performance objectives for takeoff rotation are given in table LVIII.
Determination of performance objectives for overshoot are dependent on aircraft geometry and the
recommended takeoff attitude.
After main wheel liftoff, the pilot maintains a specific flightpath angle and groundtrack. He maintains this
flightpath until the landing gear has been retracted and all transients have settled out. Some suggested
performance objectives for takeoff climbout are given in table LVIII. Determination of performance
objectives for flightpath control should consider what cues are available to the pilot. In most cases pitch
attitude is used as a substitute when flightpath angle is not an available cue. The tolerance for adequate
flightpath control should not allow a negative flightpath angle. For groundtrack control, heading angle may
be used as a substitute for deviation from runway centerline.
It would be impractical to evaluate takeoff roll with a long-look technique in flight test because the aircraft
would have to land and taxi back to the end of the runway each time. However, rotation and climb-out
could be evaluated with a long-look technique by doing touch-and-goes. On a ground-based simulator the
entire task could be done using a long-look technique. As with the landing tasks, light random turbulence
and moderate discrete gusts should be used to increase the pilot's “gain” on a ground-based simulator.
Degradation with atmospheric disturbances (4.9.1) can be evaluated on the ground simulator by
increasing the turbulence and gusts and by adding crosswinds and wind shears.
VERIFICATION LESSONS LEARNED
An important source of guidelines on the use of tracking techniques for handling qualities evaluation is
AFFTC-TD-75-1. AFFTC-TD-75-1 discusses execution and analysis of results of both air-to-air and
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APPENDIX A
air-to-ground techniques. Many of the recommendations in AFFTC-TD-75-1 are also applicable to other
closed-loop handling qualities evaluation techniques. Further discussions of the design and conduct of
handling qualities testing and the use of the Cooper-Harper scale can be found in CAL Report TB-1444F-1, NADC-85130-60, NASA TN D-5153, AlAA 89-3358, and AlAA 90-2822.
Some of the. most detailed descriptions of closed-loop evaluation tasks which have been used in the past
can be found in USAFTPS and AFFTC handling qualities test plans and flight test reports. Descriptions of
several tasks taken from USAFTPS Letter Reports can be found in AFFDL-TR-77-34 and AFFDL-TR-793126. AFFTC, handling qualities reports which contain descriptions of several closed-loop evaluation
tasks include AFFTC-TR-75-15 (the YF-16), AFFTC-TR-77-11 (the A-10), AFFTC-TR-77-23 (the YF-16
Control Configured Vehicle (CCV)), AFFTC-TR-83-45 (the Advanced Fighter Technology Integration
AFTI/F-16), and AFFTC-TR-91-29 (the F-15 STOL and Maneuver Technology Demonstrator (SIMTD)).
Another source of closed-loop evaluation tasks is the Standard Evaluation Maneuver Set (STEMS). The
results of this project are documented in WL-TR-93-3081, WL-TR-93-3082, and WL-TR-93-3083. The
main products of this project were: 1) a process to develop handling qualities evaluation maneuvers, 2)
an initial set of 20 evaluation maneuvers tested in ground simulation, and 3) guidelines to help users
select appropriate maneuvers. WL-TR-93-3081 describes the maneuver development process. WL-TR93-3082 provides descriptions of the initial set of evaluation maneuvers and a selection guide. WL-TR-933083 documents the results of the ground simulation tests of the initial maneuver set. The maneuvers
developed in this project were primarily aimed at evaluation of agility and high-angle-of-attack flying
qualities, however, there were some conventional flying qualities evaluation maneuvers as well. AIAA-933645 provides a summary of the STEMS project.
Air-to-Air Tracking
Air-to-air tracking is one of the most commonly used handling qualities evaluation techniques. Examples
of the use of this kind of task can be found in many handling qualities reports. The task descriptions and
performance objectives recommended in Verification Guidance stem largely from numerous USAFTPS
projects conducted in the mid-1970s to the early 1980s using the variable-stability NT-33A. The task
descriptions in these projects remained fairly similar throughout this period and are documented in
AFFDL-TR-77-34 and AFFDL-TR-79-3126. In the earliest of these tests the performance objectives for
Cooper-Harper ratings were undefined. The performance objectives gradually became better defined with
succeeding projects. The performance objectives suggested in Verification Guidance reflect the
objectives used in the later projects. Similar performance objectives were used on McDonnell-Douglas
and Wright Laboratory ground simulators during the development of the F-15 S/MTD (WRDC-TR-493036).
Some other target maneuvers which have been used in air-to-air tracking are a modified Lazy-Eight
maneuver, a constant-g barrel roll, and an unpredictable target maneuver. Discussions of the use of the
modified Lazy-Eight and the barrel roll maneuver can be found in AFFDL-TR-79-3126. The unpredictable
target is a target which is free to maneuver within certain restrictions. Normally. it is restricted in airspeed
(typically within ± 50 knots of test condition), altitude (typically within ± 1000 feet of test condition), load
factor, and onset rate (typically restricted to no more than .5 g/sec).
One of the conventional evaluation tasks of the STEMS project was Tracking in Power Approach. The
task was to track a target aircraft from approximately 1500 ft range in power approach configuration at
approach airspeed. The target performed gradual S-turns with periods of straight flight between turns.
Constant altitude was maintained during the maneuver. The evaluation pilot selected specific aim points
on the target and tracked them during the maneuver. In the simulation, different target profiles were
required for different Classes of aircraft. For fighter aircraft, the target performed a 30° heading change
vary 20 seconds. For transport aircraft, a 15° heading change was performed every 15 seconds. Desired
performance was to keep the pipper within ±5 mils of the aim point for 50% of the task and within ±25 mils
for the remainder of the task, with no PIO. Adequate performance was to keep the pipper within ±5
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MIL-STD-1797A
APPENDIX A
mils of the aim point for 10% of the task and within ±25 mils for the remainder of the task. This maneuver
can be performed at a safe altitude before attempting precision landings.
HUD tracking tasks have been used in a number of handling qualities research programs. Some example
step-and-ramp tracking sequences are shown on figure 273. The sequences shown in a) and b) are pitch
tracking sequences that were used on the NT-33A in USAFTPS projects (USAFTPS Report 8213-4). The
two sequences shown in c) are a combined pitch and roll tracking sequence used on the NT-33A and on
Calspan's variable-stability LearJet in many recent projects (Calspan Report No. 7738-24). Another type
of sequence in use is a sum-of-sines. This is a frequency-based function driven by an equation such as:
n
Θc = K
∑A
i
sin(ωi t + φi )
i =1
where
Θc =
the commanded pitch attitude (or bank angle)
K
=
the task gain
n
=
the number of sine waves
Ai
=
the amplitude of each sine wave
ωi
=
the frequency of each sine wave
φi
=
the phase of each sine wave
Such a function was used in a project on the LearJet (AFFTC-TLR-93-41) using 13 sine waves evenly
spaced in frequency between 0.1 and 30 rad/sec. The HUD symbology usually used in HUD tracking
tasks are tracking bars, but with the computational power and electronic displays available today, it is
worth considering special flight test software to provide a more definitive target.
One advantage of HUD tracking tasks is that, if the HUD update rate and HUD dynamic characteristics
are duplicated in the simulators, the task itself can be identical between ground simulation, in-flight
simulation, and flight test, providing a greater degree of commonality between these three stages of
evaluation. Bear in mind, though, that HUD dynamic characteristics will affect handling qualities more in
these tasks than in tasks such as formation flying or VFR landings.
Air-to-Ground Tracking
Previous flight test programs which have used air-to-ground tracking techniques include the A-10
(AFFTC-TR-77-11 and Brandeau, AFFDL-TR-78-171). YF-16 CCV (AFFTC-TR-77-23), and AFTI/F-1 6
(AFFTC-TR-83-45). A very promising system for an air-to-ground tracking task called GRATE (Ground
Attack Targeting Equipment) was developed and tested by the Deutsche Forschungs-und
Versuchsanstalt fur Luft-und Raumfahrt (DFVLR) in the 1980s (Koehler, NASA CP 2428 and Koehler,
AGARD-CP-452). This task used a pattern of lights on the ground as a target. The pilot acquires and
tracks each light in turn as the lights are illuminated in a specific sequence. The Germans evaluated this
task with an AlphaJet with great success. The task was also subsequently used successfully on a groundbased simulator (Biezad, AFWAL-TR46-3093). In 1987 the Dryden Flight Research Center (DFRC)
developed a derivative system known as the Adaptable Target Lighting Array System (ATLAS). This
system has been tested with the NT-33 (USAFTPS-TR-88A-TM 1), the X-29 (NASA TM 101700), and the
F-15 S/MTD. An example of a typical ATLAS array is shown on figure 274. The ATLAS system is
currently operational at DFRC at Edwards AFB.
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APPENDIX A
FIGURE 273. Example step-and-ramp HUD tracking sequences.
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APPENDIX A
FIGURE 273. Example step-and-ramp HUD tracking sequences - continued.
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MIL-STD-1797A
APPENDIX A
FIGURE 274. Typical ATLAS light pattern (from NASA TM 101700).
Close Formation Task
Formation flying tasks were used in the evaluations of the YF-16 CCV (AFFTC-TR-77-23), the AFTI/F-16
(AFFTC-TR-83-45), and the F-15 S/MTD (AFFTC-TR-91-29 and WL-TR-92-3027). The YF-16 CCV
evaluation included station keeping and position changes during straight and level flight and in “mild to
moderate" lazy-eight maneuvers. Vertical and horizontal position changes of 500 feet and 50 feet were
made.
AFTI/F-16 formation tasks were done in close trail and wingtip positions. The following descriptions are
taken from AFFTC-TR-83-45.
“The close trail evaluations were conducted with the AFTI/F-16 in a position ten feet below the
lead aircraft with zero to ten feet nose-to-tail separation. The lead aircraft flew between 250 to
300 KCAS while making gentle maneuvers. This maneuvering included bank angle variations up
to 30 degrees and gentle climbs and descents."
"The close formation evaluations were conducted with the AFTI/F-16 in a tight position off the
wing of the lead aircraft. After a reasonable buildup, the AFTI/F-16 pilot flew as close to the lead
aircraft as was comfortable. The lead aircraft flew either in level flight or through a series of lazy
eight maneuvers. Bank angle changes ranged between ±90 degrees, pitch attitude ranged
between ±45 degrees, airspeed between 200 and 500 KCAS and load factor between one and
five g's. In this task the pilots attempted to maintain a precise position relative to the lead aircraft
using fingertip formation techniques. The AFTI/F-16 would intentionally make lateral and vertical
deviations in order to evaluate the aircraft's ability to return to the nominal position."
In the F-15 S/MTD program a formation task was used to evaluate PIO tendencies in power approach.
The pilot attempted to maintain formation with an A-37 while the A-37 performed random 0.25-g step
inputs (three to five foot vertical variations).
Aerial Refueling
Boom tracking with a HUD pipper or a waterline symbol has been done in several programs at AFFTC. In
every case, however, it was done using the HQDT technique (described in 5.1.11.6 Verification
Guidance), where the objective was zero pipper error. At this time there is no data to support the
suggested performance objectives for this task other than the fact that these objectives have been
successfully applied in other kinds of tracking tasks.
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APPENDIX A
Boom tracking using visual alignment with the tanker was proposed as a handling qualities task by the
4950th Test Wing in a flight test project to develop an aerial refueling evaluation task for Class III aircraft.
The tanker aircraft in their project was a KC-135. The visual desired zone for their task was between the
rivet lines on the bottom of the KC-135. The adequate zone was the edges of the fuselage. A diagram of
these zones is shown in Figure 275. In the test they used two different evaluation aircraft: a C-135 and a
C-18. The performance objective was the cumulative time the evaluation pilot could keep the boom
nozzle aligned within the desired zone during two minutes of tracking. Four different levels of desired
performance were tested: 30 seconds, 45 seconds, 60 seconds, and 75 seconds. Adequate performance
was defined as keeping the nozzle within the adequate zone for the entire 2 minutes. Unfortunately, the
pilots considered both aircraft Level 2 for this task because of the amount of compensation required. The
results for all four levels of desired performance gave Level 2 C-H ratings for both aircraft, thus the C-H
ratings give no indication of which is the best value to use for desired performance. The results do lend
credence to their choice of adequate criteria, but this was not a variable in the test. Nevertheless, based
on the performance achieved with both aircraft throughout the project, the pilot consensus was that 60
seconds in the defined desired zone (or about 50% of the tracking time) was “both attainable and
realistic" and that the task was demanding enough to expose undesirable handling qualities. This project
was documented in 4950-FTR-93-05.
FIGURE 275. Adequate and desired performance for 4950th
boom tracking task (from 4950-FTR-9345).
The STEMS project also tested a boom tracking task. The task was to track the refueling probe of a
tanker from the pre-contact position. The evaluation pilot can track a steady probe, periodically changing
aim points on the boom (such as the boom wingtips), or the boom operator can make small random
horizontal and vertical movements with the boom to create tracking errors. Desired performance in the
STEMS project was to maintain the aim point within a 30-mil radius of the pipper for at least 50% of the
task, with no objectionable PlOs. Adequate performance was to maintain the aim point within a 50-mil
radius of the pipper for at least 50% of the task.
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MIL-STD-1797A
APPENDIX A
Probe-and-drogue refueling was used as an evaluation task in an experiment with the NT-33A
documented in AFFDL-TR-74-9. Standard probe-and-drogue refueling procedures were used. However,
adequate and desired performance objectives were not explicitly defined for this project. The performance
objectives suggested in Verification Guidance are taken from a USAFTPS study of response-types, for
probe-and-drogue refueling performed with the NT-33A in October 1993.
A position-keeping evaluation task has been used extensively for tanker evaluations including the S-3
with a buddy store, the KC-10, the KC-130, and the KC-135. A detailed description of this test technique
and lessons learned from these test programs is provided in AGARD CP-519.
Offset Precision Landing
This is another task which is so widely used for handling qualities evaluations that descriptions of it can
be found in many handling qualities reports, particularly those that deal with approach and landing. The
suggested performance objectives for the approach and touchdown phases are taken largely from
Calspan experience with the NT-33A and the Total In-Flight Simulator (TIFS). Discussions of the use of
this task can be found in NASA CR 172491, NASA CR 178188, AFWAL-TR-81-3118, and AIAA 93-3816.
Some data on the size of the suggested landing zone was provided in a USAFTPS project which studied
the effect of different performance objectives on touchdown C-H ratings with three different Class IV
aircraft: the F-15D, the F-16D, and the F/A-18B. The experiment looked at three different desired landing
zones: 25 feet wide by 200 feet long. 50 feet wide by 400 feet long, and 75 feet wide by 600 feet long. CH ratings for each aircraft and each landing zone were compared with the pilots' qualitative appraisal of
each aircraft. For the 75x600 foot zone, all three aircraft received basically Level 1 C-H ratings. For the
50x400 foot zone, the F/A-18 received basically Level 1 C-H ratings, the F-15 had borderline Level
1/Level 2 C-H ratings, and the F-1 6 got basically Level 2 C-H ratings. For the 25x200 foot zone, all three
aircraft received basically Level 2 C-H ratings. The project report found the results from the 50x400 foot
zone to be most representative of the pilots' qualitative opinions of each aircraft.
Additional data on the size of the suggested landing zone was provided in a project by the 4950th Test
Wing which studied the effect of different performance objectives on touchdown C-H ratings with three
different Class III aircraft: the C-141A, the C-135A/E, and the C-18B. The experiment looked at four
different desired landing zones: 20 feet wide by 200 feet long, 40 feet wide by 400 feet long, 60 feet wide
by 800 feet long, and 80 feet wide by 1000 feet long. C-H ratings for each aircraft and each landing zone
were compared with the pilots' qualitative appraisal of each aircraft. All three aircraft were considered
Level 2 for the landing task. The final report (4950-FTR-93-05) recommended the 40x400 foot zone. For
this zone, desired performance was met 6 out of 12 times, and the pilot ratings comprised two C-H ratings
of 3 and four C-H ratings of 4. Interestingly, this suggested landing zone is very similar to the one
recommended by USAFTPS for Class IV aircraft.
The performance objectives for STOL aircraft are taken from experience on the F-15 S/MTD program.
Discussions of the tasks used in the S/MTD program can be found in AFWAL-TM-87-180-FIGC. WRDCTR-89-3036, AFFTC-TR-91-29, and WL-TR-92-3027.
Takeoff
The takeoff task described in Verification Guidance was taken from the F-15 S/MTD program.
This task was only used as an evaluation task on the McDonnell-Douglas and Wright Laboratory
simulators. It was not used as an evaluation task in the flight test program. Descriptions and
results of the use of this task can be found in AFWAL-TM-87-180-FIGC and in WRDC-TR-893036.
Supersedes page 109 of MIL-STD-1797A
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MIL-STD-1797A
APPENDIX A
4.1.7 Aircraft Failure States.
The contractor shall define and tabulate all Aircraft Failure States which can affect flying qualities. Aircraft
Failure States consist of Aircraft Normal States modified by one or more malfunctions in aircraft
components or systems; for example, a discrepancy between a selected configuration and an actual
configuration. Those malfunctions that result in center-of-gravity positions outside the center-of-gravity
envelope defined in 4.1.1 shall be included. Each mode of failure shall be considered in all subsequent
Flight Phases.
REQUIREMENT RATIONALE (4.1.7)
This tabulation is the starting point for a failure modes and effects analysis, which is necessary in a
complex aircraft to assure flying qualities adequate for mission effectiveness and flight safety.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.6.2.
Because of the exhaustive work often involved and low confidence in the failure probability calculations,
there is a tendency for the procuring activity to substitute a prior list of specific failures. If a design is far
enough along and not excessively complex, such an approach can work. See the guidance for 4.1.7.1.
However, generally comprehensive reliability analyses will be required anyway.
Whether the approach to failure effects on flying qualities is probabilistic, generic or a combination, failure
possibilities of the specific aircraft must be catalogued thoroughly enough to assure adequate mission
effectiveness and flight safety.
REQUIREMENT LESSONS LEARNED
There is more to determining Failure States than just considering each component failure in turn. Two
other types of effects must be considered. First, failure of one component in a certain mode may itself
induce other failures in the system, so failure propagation must be investigated. Second, one event may
cause loss of more than one part of the system or can affect all channels: a broken bracket, a single
crack, a fire, an electrical short, inadequate ground checkout, etc. The insidious nature of possible
troubles emphasizes the need for caution in design applications.
5.1.7 Aircraft Failure States - verification.
The contractor shall furnish the required data in accordance with the Contract Data
Requirements List.
VERIFICATION RATIONALE (5.1.7)
Definition of aircraft failure states is basic to the application of the flying qualities requirements.
VERIFICATION GUIDANCE
Generally, compliance will amount to identifying pertinent hems from the list required by the reliability
specification, and checking for completeness. Although the task may seem formidable, the alternative to a
thorough review is a certainty that something important will be overlooked.
VERIFICATION LESSONS LEARNED
REPRINTED WITHOUT CHANGE
110
MIL-STD-1797A
APPENDIX A
TABLE X. Control-margin increments.
Flying Quality
Stabilization
Turbulence
∆δ FQ / ∆n c = 57.3 CAP ′ / M δ deg/ g (for Teff ≤ 0.5 )
∆δ stab / ∆nc = 57.3
g 1 Tsp1 ⋅ 1 Tsp 2
⋅
deg/ g (linear, 2 DOF)
Uo M δ ⋅ 1 TΘ 2
σ δ / σ w fn of Mw , Mδ , ωsp cl , ζ sp cl , structural m
- most severe at low q
- 3 σ δ and σ w for severe turbulence recommended
Sensor Noise
σδ / σs fn of K s , K F , ωs , 1 Ta , ωsp cl , ζ sp cl , ω2sp ol
Flying Quality
δ& FQ / nc = 57.3 CAP / (M δ ⋅ T eff ) for desired CAP
Stabilization
δ& stab / nc < δ& FQ / nc if FCS stability margins OK &
1
> ωc
Teff
δ& stab / nc fn of 1 Teff , 1 Tsp 2 , ωsp cl , ζ sp cl if FCS stability margins OK &
Turbulence
1
> ωc
Teff
σδ& / σ w fn of 1 Ta , ωspcl , ζ spcl , Mδ
- most severe at low q
- 3σδ& recommended for control margin
Sensor Noise
(
σ δ& / σ s = K s , K F ⋅ fn ω s , 1 Ta and, for low ω spcl : ω 2spol , ω spcl , ζ spcl
)
- These parameters are not all independent
- 3σδ& recommended for control margin
∆n c is the commanded increment of normal acceleration
1 T2 is the unstable pole of the transfer function (negative, 1/sec)
ω2sp cl is the 2-deg-of-freedom product of the poles, 1/sec2
ω
sp cl
and ζ sp cl are the closed-loop frequency and damping ratio of the short period mode
CAP is q& o / ∆nοο , CAP′ is q& max / ∆ οο
ωs is the sensor bandwidth
Ks, KF are the sensor and forward-loop gains
σs , σ w are the rms intensities of sensor noise and vertical gusts
ωc is the crossover frequency of the δ& / nc transfer function
Teff is the effective time constant of command-path plus forward-path control-loop elements (such as
prefilters and actuators)
Ta is the time constant of actuator ram
Supersedes page 149 of MIL-STD-1797A
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MIL-STD-1797A
APPENDIX A
FIGURE 11. Control margin requirements.
REPRINTED WITHOUT CHANGE
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MIL-STD-1797A
APPENDIX A
4.1.11.6 Pilot-in-the-loop oscillations (PIO).
There shall be no tendency for pilot-in-the-loop oscillations, that is, sustained or uncontrollable oscillations
resulting from efforts of the pilot to control the aircraft. More specific requirements are in 4.2.1.2, 4.2.2,
4.5.2, and 4.6.3.
REQUIREMENT RATIONALE (4.1.11.6)
This general qualitative requirement, applicable to all axes, covers those axes of control for which there is
no data base for more specific requirements.
REQUIREMENT GUIDANCE
The applicable MIL-F-8785C requirements are paragraphs 3.2.2.3 and 3.3.3.
PIOs were a consideration in setting the boundaries of 4.2.1.2 and 4.5.1.3 through 4.5.1.5.
REQUIREMENT LESSONS LEARNED
Likely causes are equivalent time delay control system friction, inappropriately-located zeros of aircraft
transfer functions, or nonlinearities such as rating limiting, hysteresis, and abrupt control system gain
changes. See the discussion under 4.2.2. NORAIR Rpt No. NOR-64-143 discusses a number of possible
PIO mechanisms.
5.1.11.6 Pilot-in-the-loop oscillations (PIO) – verification.
Verification shall be by analysis, simulation, and flight test. Final verification shall be by demonstration in
the following tasks: __________________.
VERIFICATION RATIONALE
The need to use high-gain, closed-loop tasks to evaluate handling qualities, in addition to comparison
with open-loop requirements if fully discussed in the Verification Rationale of 5.1.6.1 through 5.1.6.3. An
additional reason, if any more are needed, is that most of the open-loop requirements assume a linear
system. Pilot evaluation in high-gain, closed-loop tasks is at this time the best evaluation of the effects of
nonlinearities. This is particularly important in the evaluation of PIO tendencies because nonlinearities,
such as rating limiting, hysteresis, abrupt gain changes, and aerodynamic nonlinearities, are some of the
common causes of PIO.
VERIFICATION GUIDANCE
Pilot-vehicle analysis in the manner described in the discussion of the cited paragraphs should help in the
design stage. Ground-based simulation may or may not show up any PIO tendencies. Flight evaluation in
variable-stability aircraft is a valuable tool. Final determination will come from flight test of the actual
vehicle.
The recommended tasks to demonstrate compliance with this requirement are the tasks described in
Verification Guidance of 5.1.6.1 through 5.1.6.3 using the HQDT technique. AFFTC makes a distinction
between HQDT and “operational” closed-loop evaluation tasks. The key element of the HQDT technique
is that the pilot must attempt to totally eliminate any error in the performance are not defined and
aggressive control strategy that he can. Adequate and desired performance are not defined and CooperHarper ratings are not recommended. The reason for this is that, in the “operational” tasks, definition of
adequate and desired performance encourages the pilot to adopt a control strategy which best meets
these performance objectives. In the case of a PIO-prone airplane, attempting to totally eliminate any
deviation may induce oscillations which reduce his performance, but by accepting small errors (reducing
his gain) the pilot may be able to avoid these oscillations and still meet the performance objectives
(which, by their definition, allow such a tactic). The HQDT technique does not allow the pilot to do this,
thus exposing any possible handling qualities deficiencies. HQDT could be considered a “stress test” of
handling qualities. For this reason, the HQDT technique is considered the best test of PIO tendencies.
HQDT is not exclusively a PIO evaluation technique. It is a general handling qualities evaluation
technique. It is discussed in more detail here in the PIO requirement because it is a better PIO
Supersedes pages 151-152 of
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MIL-STD-1797A
APPENDIX A
”operational” technique uses performance objectives more representative of operational use, and the C-H
ratings provide a quantitative measure of flying qualities which can be related to the required Levels.
Therefore, use of both techniques is recommended in the flight test evaluation, as well as parameter
identification techniques and capture tasks. As mentioned in the Verification Rationale of 5.1.6.1 through
5.1.6.3, the recommended parts of the handling qualities evaluation are: 1) steps, doublets, and
frequency sweeps for parameter identification and comparison to open-loop requirements, 2) capture
tasks for pilot familiarization with aircraft dynamic response and evaluation of gross acquisition, 3) HQDT
for initial handling qualities and PIO evaluation (HQDT may also provide good inputs for frequencydomain analysis), and 4) “operational” tasks for handling qualities evaluation with C-H ratings.
The PIO tendency classification scale shown on figure 12 has been developed specifically for evaluation
of PIO tendencies. It can be used with either the HQDT or the “operational” techniques. Comparing the
PIO rating descriptions with descriptions of Levels of flying qualities, a rough approximation would be:
PIO ratings of 1 or 2 would be Level 1, PIO ratings 3 or 4 would be Level 2, and a PIO rating of 5 would
be Level 3. A PIO rating of 6 would be extremely dangerous.
FIGURE 12. PIO tendency classification from AFWAL-TR-81-3118.
Supersedes pages 151-152 of
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MIL-STD-1797A
APPENDIX A
Tom Twisdale provides some guidance on possible HQDT tasks:
Probably any test maneuver that allows the evaluation pilot to aggressively and
assiduously track a precision aim point is a suitable HQDT test maneuver. In HQDT
testing, the test maneuver is not nearly as important as the piloting technique. It is the
piloting technique that increases the evaluation pilot’s bandwidth and makes possible a
good handling qualities evaluation. For this reason there is no exclusive catalog of HQDT
maneuvers. The ones discussed below have worked well, but others, perhaps better
suited to a particular airplane, may be invented as the need arises.
Air-to-Air HQDT
Air-to-Air HQDT involves tracking a precision aimpoint on a target airplane while using a
fixed, or non-computing gunsight. There are three main variations of air-to-air HQDT:
constant load factor HQDT at a constant range of about 1500 feet; wind-up turn HQDT at
a constant range of about 1500 feet; and HQDT while closing on the target.
The purpose of a constant load factor air-to-air HQDT maneuver is to evaluate handling
qualities at a specific angle of attack. The maneuver begins with the test airplane
positioned 1500 feet behind and offset above, or below, or to the inside of the target. The
offset position is helpful in avoiding jetwake encounters. At the evaluation pilot’s signal
the target pilot rolls smoothly into a turn and slowly increases load factor until the test
load factor is attained. A g onset rate of two seconds or so per g is satisfactory. When the
test load factor has been attained the target pilot calls “on condition” and maintains the
turn and the test conditions for the specified period of time, which will depend on the test
and analysis objectives. During the load factor build-up the evaluation pilot turns on the
airborne instrumentation system and positions the target airplane 50 mils or so from the
pipper or aiming index at a clock position of 1:30, 4:30, 7:30, or 10:30. After the target
pilot calls “on condition” the evaluation pilot calls “tracking” and drives the pipper toward
the precision aim point to initiate the evaluation. The evaluation pilot continues to track
while using the HQDT piloting technique, until the target pilot or other aircrew or the
control room calls ”time”. However the maneuver is not concluded until the evaluation
pilot calls “end tracking”. At that time the target pilot rolls out of the turn.
The constant load factor air-to-air HQDT maneuver may be a constant turn to the left or
right, or turn reversals may be included. When reversals are included they should be
performed at constant load factor. The evaluation pilot continues to track the precision
aim point throughout the reversal, always using the HQDT piloting technique.
The evaluation pilot should maintain a 1500-foot separation from the target airplane.
Variations of a few hundred feet either way are permissible, but range to the target
should not be allowed to exceed 2000 feet. Range may be determined stadiometrically
with adequate accuracy.
The purpose of a wind-up turn air-to-air HQDT maneuver is to quickly explore handling
qualities across a range of angle of attack. The maneuver gets under way when the
target pilot establishes the test conditions and calls “on condition”. The evaluation pilot
positions the test airplane 1500 feet behind and offset above, or below, or to the inside of
the target. The offset position is helpful in avoiding jetwake encounters. The evaluation
pilot turns on the airborne instrumentation system and positions the target airplane 50
mils or so from the pipper or aiming index at a clock position of 1:30, 4:30, 7:30, or 10:30.
The evaluation pilot then signals at the target pilot to begin the maneuver. The target pilot
rolls smoothly into a turn and slowly increases load factor at a g onset rate of five or six
seconds per g. As the target airplane begins rolling into the wind-up turn, the evaluation
pilot calls “tracking” and drives the pipper toward the precision aim point to initiate the
evaluation. The evaluation pilot continues to track while using the HQDT technique, until
the target pilot attains the target load factor and calls “target g”. The target load factor is
maintained until the evaluation pilot calls “end tracking”. At that time the target pilot may
unload and roll out of the turn.
NEW PAGE
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MIL-STD-1797A
APPENDIX A
The evaluation pilot should maintain a 1500-foot separation from the target airplane.
Variations of a few hundred feet either way are permissible, but range to the target
should not be allowed to exceed 2000 feet. Range may be determined stadiometrically
with adequate accuracy.
In HQDT with closure, the evaluation pilot slowly closes on the target airplane while
tracking. The purpose of the closing HQDT maneuver is to help the evaluation pilot
distinguish attitude dynamics from normal and lateral acceleration dynamics. Attitude
dynamics are evident at any tracking range, but translation caused by normal and lateral
acceleration become more noticeable as the evaluation pilot closes on the target.
In a closing HQDT maneuver the target airplane may either fly straight and level,
maneuver gently in pitch and roll, or perform a constant load factor turn. Gently
maneuvering or a constant load factor turn is often preferred because it helps to increase
the evaluation pilot’s bandwidth. In all other respects the closing maneuver is similar to a
constant load factor or wind-up tracking turn.
The closing HQDT maneuver can begin once the target pilot has established the test
conditions and calls "on condition". The evaluation pilot positions the test airplane 1500
feet behind and above, below, or to the inside of the target; turns on the airborne
instrumentation system; and positions the target airplane 50 mils, or so from the pipper or
aiming index at a clock position of 1: 30, 4:30, 7:30, or 10:30. The evaluation pilot then
signals the target pilot to begin the maneuver. The target pilot flies straight and level: or
begins to maneuver gently and randomly in pitch and roll; or performs a constant load
factor turn. The evaluation pilot calls "tracking" and drives the pipper toward the precision
aim point to initiate the evaluation. The evaluation pilot continues to track, using the
HQDT technique, while slowly dosing on the target airplane. The rate of closure will
depend on the desired tracking time (which will depend on the test and analysis
objectives). The evaluation pilot may find it easier to control the rate of closure 6 the
control room or the target pilot or other aircrew announce the elapsed time in five second
increments. At the end of the specified tracking time, the target pilot or other aircrew or
the control room calls "time". However the maneuver is not concluded until the evaluation
pilot calls "end tracking".
Power Approach HQDT
Power approach HQDT is air-to-air HQDT performed with the test airplane configured for
power approach. This maneuver is designed to evaluate approach and landing handling
qualities at a safe altitude (10,000 to 15,000 feet), rather than a few feet above the
ground during a real landing. Power approach HQDT maybe flown with or without
closure, however closure is a desirable feature because it helps the evaluation pilot
distinguish between attitude and translation dynamics.
The target airplane may either fly straight and level or maneuver gently in pitch and roll.
Maneuvering gently is often preferred because it helps to increase the evaluation pilot's
bandwidth. In all other respects the power approach HQDT maneuver is similar to a
closing HQDT maneuver.
Closure during the maneuver is useful for distinguishing attitude dynamics from normal
and lateral acceleration dynamics. Attitude dynamics are evident at any tracking range,
but translation caused by normal and lateral acceleration become more noticeable as the
evaluation pilot closes on the target.
Jet-wake encounters are a frequent source of difficulty during power approach HQDT
testing. Simple geometry, together with a maneuvering target airplane, make jet-wake
encounters difficult to avoid. The slow speeds introduce the risk that a jet-wake encounter
will precipitate a stall or departure, although this has never occurred. There are two
solutions to the problem of jet-wake encounters. One is to use a small propeller-driven
airplane as a target. Excellent candidates are the T-34C or Beechcraft Bonanza, or
similar airplanes.
152b
MIL-STD-1797A
APPENDIX A
These airplanes can easily match the slowest speeds of most military airplanes, and they
produce very little propwash. The second solution is to use a target that is programmed
into a flight test head-up display, similar to the head-up display used on the Calspan NT33.
Air-to-Ground HQDT
Air-to-ground HQDT involves tracking a precision aimpoint on the ground with a fixed, or
non-computing gunsight. Shallow or steep dive angles may be used. Shallow dive angles
approximate strafing attack profiles and steeper angles approximate ballistic weapons
delivery profiles.
The evaluation pilot trims the airplane at the specified dive entry altitude and airspeed,
turns on the airborne instrumentation system, calls "on condition", and rolls or pitches to
the specified dive angle. When the outer ring of the gunsight reticle crosses the precision
aim point, the evaluation pilot calls "tracking" and commences to track the precision aim
point using the HQDT piloting technique. The evaluation pilot continues to track until the
recovery altitude is reached, then calls "end tracking" and recovers from the dive.
A useful variation on the basic maneuver is to track two or more precision aim points,
instead of one. For example, precision aim points may be positioned at each apex of an
imaginary isosceles triangle laid out on the ground. This triangle has a base of 100 feet
and a height of 375 feet (for 15 degree dive angles) or a height of 100 feet (for 45 degree
dive angles) . During the maneuver the evaluation pilot randomly switches from one
precision aim point to another, perhaps at a signal from the control room.
Boom Tracking HQDT
In boom tracking, the evaluation pilot tracks the nozzle on an aerial refueling boom. This
maneuver is designed to explore aerial refueling handling qualities without the risk of
close proximity to a tanker and a refueling boom.
The tanker airplane establishes the test conditions of Mach number (or airspeed) and
altitude and maintains them during the test maneuver. The boom operator positions the
refueling boom at zero degrees of azimuth and a midrange elevation angle. When the
test conditions have been established the tanker pilot or the boom operator call "on
condition". The evaluation pilot moves the test airplane into position a short distance
behind the nozzle (20 to 50 feet) and positions the nozzle about 50 mils from the pipper
or aiming index at a clock position of 1:30, 4:30, 7:30, or 10:30. To begin the maneuver,
the evaluation pilot turns on the airborne instrumentation system, calls "tracking", and
drives the pipper toward the nozzle. The evaluation pilot continues to track the nozzle,
using the HQDT piloting technique, while the boom operator randomly maneuvers the
refueling boom in azimuth and elevation. The boom motion should be a combination of
gentle and abrupt changes in rate and position. After the specified period of tracking time
(which will depend on the test and analysis objectives) has elapsed, the control room or
another crew member calls "time". The maneuver is not concluded, however, until the
evaluation pilot calls "end tracking".
Formation HQDT
In formation HQDT, the evaluation pilot attempts to maintain a precisely defined position
relative to the lead airplane during a series of gentle maneuvers. Properly done,
formation HQDT can highlight for the evaluation pilot the vertical and lateral translation
dynamics of the test airplane. This maneuver is also useful for evaluating the throttle
response of the airplane. Care must be taken not to force the evaluation pilot to fly too
close to the lead airplane. Close proximity can increase bandwidth, but too close
proximity can reduce it. As the separation between airplanes narrows, good and prudent
pilots will reduce their bandwidth to reduce the risk of collision.
152c
MIL-STD-1797A
APPENDIX A
Attention to flying qualities per se during flight control design will take care of many potential problems.
PlOs may occur early in the aircraft life as on the YF-16 high speed taxi test that got airborne before its
first flight, or later in service, as with the T-38 as more pilots got to fly it. If PIO is not found readily, it
should be sought during the flight test program.
152d
MIL-STD-1797A
APPENDIX A
4.2.2 Pilot-in-the-loop pitch oscillations.
There shall be no tendency for sustained or uncontrollable pitch oscillations resulting from efforts of the
pilot to control the aircraft. The phase angle of the pitch attitude frequency response to pitch stick force at
the criterion frequency, ωc, shall be greater than or equal to ______. If this phase angle is less than
______, the phase parameter of normal acceleration at the pilot's station, φ, at the same criterion
frequency, shall be greater than or equal to ______. Furthermore, the requirements of 4.2.1.2, 4.2.8.1,
4.2.8.2, and 4.2.8.4 must be met.
REQUIREMENT RATIONALE (4.2.2)
The purpose of this requirement is to insure that abrupt maneuvers or aggressive tracking behavior will
not result in instabilities of the closed-loop pilot/aircraft system. Any such tendency will degrade or even
destroy mission effectiveness and likely will be dangerous.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.2.2.3.
Recommended values:
The recommended minimum phase angle of the pitch attitude frequency response to the pitch stick force
at the criterion frequency, ωc, is -180°.
The value of ωc is determined as follows:
ωc (0.24 rad-oct/dB-sec) S + 6.0 rad/sec
where S is the average slope of
Θ(s)
in dB/oct over the interval from 1 to 6 rad/sec.
Fes (s)
If the phase angle of pitch attitude frequency response to pitch stick force is less than -160°, then the
recommended minimum value of the phase parameter of normal acceleration at the pilot's station, φ, at ωc
is -180°.
The phase parameter of normal acceleration at the pilot's station, φ, is defined by:
φ(ωc ) = ∠
( )
−
Fes (jω )
a z p jω c
(14.3 deg-sec/rad) ωc
c
where a z p is normal acceleration at the pilot's station.
A related requirement in 4.2.8.2. Also, see 4.1.11.6 for a general PIO requirement. The qualitative
requirement of MIL-F-8785C is generalized in view of uncertainties in the state-of-the-art of flight control
system design, a tacit recognition of the complexity of the PIO problem; no detailed specification is, at this
time, a guarantee against building a PIO-prone airframe/flight-control-system combination.
The requirement precludes PIO, PIO tendencies or general handling qualities deficiencies resulting from
inadequate pilot-vehicle closed-loop gain and phase margins. PIO has occurred in the T-38A. A4D, and
YF-12 due to abrupt amplitude-dependent changes in aircraft dynamic response to pilot control inputs.
These effects can be of mechanical origin, e.g. bobweights coupled with static friction, or due to
saturation of elements within the control system, or due to compensation added to the automatic control
system. Other known sources are short-period dynamics (e.g. large ωSP TΘ ), feel system phasing (e.g.
2
effective bobweight location not far enough forward), and sensitive control force and motion gradients.
AFFDL-TR-69-72 and Norair Rpt NOR-64-143 can furnish some insight.
The requirement above is popularly known as the Smith-Geddes PIO criteria. It was proposed in its
original form in AFFDL-TR-77-57. It was more fully developed as a general longitudinal response
requirement in AFFDL-TR-78-154, and further developed and extended to the lateral-directional axis in
Supersedes pages 267-274 of
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MIL-STD-1797A
APPENDIX A
AFWAL-TR-81-3090. No specific method for determining S was required in AFFDL-TR-78-154. Smith
originally recommended a range of 2 to 6 rad/sec and computed S as the average of three linear
approximations of slopes between pairs of points one octave apart, i.e.:
S =

1 Θ
(4 j) − Θ (2 j) + Θ (5 j) − Θ (2.5 j) + Θ (6 j) − Θ (3 j) 

3  Fes
Fes
Fes
Fes
Fes
Fes

In AFWAL-TR-81-30990 the frequency range was extended to 1 to 6 rad/sec, and a similar formula was
used to compute S using five slopes instead of three, i.e.:
 Θ

(2 j) − Θ (1j) + Θ (3 j) − Θ (1.5 j) + Θ (4 j)


Fes
Fes
Fes
Fes
1  Fes

S= 

3
Θ
Θ
Θ
Θ
Θ

(
(
(
(
(
−
2 j) +
5 j) −
2. 5 j ) +
6 j) −
3 j) 


Fes
Fes
Fes
Fes
Fes


This modification brought predictions of ωc more in line with observations based on Landing Approach
Higher Order System (LAHOS) data. AFFDL-TR-78-154 advises using a consistent method to calculate
S , even when the slope of
Θ(s )
varies considerably in the 1 to 6 rad/sec range.
Fes (s)
The criterion frequency, ωc, is an approximation of the crossover frequency of the pilot-vehicle system
during pitch attitude tracking. This approximation is based on data from AFFDL-TR-65-15 shown on figure
276. This figure shows crossover frequency as a function of forcing function bandwidth for different
controlled elements. The equation for ωc was derived from this data as shown on figure 277, taken from
AFFDL-TR-78-154. This equation was altered slightly in AFWAL-TR-81-3090 to the form recommended in
Requirement Guidance. This modification was made in conjunction with the modifications in the
calculation of S to better fit the LAHOS data and F-15 CAS-off, supersonic PIO experiences.
The third parameter in this requirement, φ(ω c ) , is a phase parameter associated with the normal
acceleration sensed at the pilot's station. It consists of the aircraft phase angle of normal acceleration
frequency response plus a phase angle due to an assumed pilot response delay at the pitch attitude
criterion frequency, ωc. This parameter becomes important if there is too much phase lag in the pitch
attitude response to stick force; thus the condition that this parameter be considered when the phase
angle of the pitch attitude response to stick force is less than -160°. This is a fundamental element of the
Smith-Geddes, originally derived in AFFDL-TR-77-57, and was included in its original form in MIL-F8785C and MIL-STD-1797. In the original form, φ was evaluated at ωR, where ωR was defined as any
frequency within the range of 1 to 10 rad/sec at which lightly damped (resonant) oscillations in pitch
attitude could result from turbulence inputs or from piloted control of the aircraft when used in the
intended manner. In AFFDL-TR-78-154, ωR was replaced by ωc as defined above. The concept behind
this part of the requirement is that, if the pitch attitude (θ) loop is resonant at ωc, then the pilot may
attempt to control normal acceleration, a zp , instead of θ. The aircraft will be PIO prone if there is too
much phase lag in this response. The criteria for this requirement are based largely on correlation with
the Neal-Smith data base (AFFDL-TR-70-74).
The statement that requirements 4.2.1.2, 4.2.2, 4.2.8.1, 4.2.8.2. and 4.2.8.4 must also be met would
seem to be redundant, since these are already requirements. However, recent history would seem to
indicate that, because the term PIO does not appear in these requirements, the importance of these
requirements in precluding PIO is not appreciated. Many recent PIO incidents can be traced directly to
problems addressed by these requirements. Therefore, these requirements are repeated here in the PIO
requirement to insure that their significance in precluding PIO tendencies is understood.
Supersedes pages 267-274 of
MIL-STD-1797A
268
MIL-STD-1797A
APPENDIX A
FIGURE 276. Variation of Crossover Frequency with Pitch
Attitude Dynamics (from AFFDL-TR-65-15).
FIGURE 277. Specification of the Criterion Frequency.
Supersedes pages 267-274 of
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269
MIL-STD-1797A
APPENDIX A
Gibson’s views of PlOs are taken directly from ICAS-86-5.3.4:
High order characteristics are associated with pilot-vehicle closed-loop handling problems
or PlO. As this term has been used to describe low order problems, the differences
should be clearly understood. The abrupt pitch bobble type is discontinuous, consisting of
repeated tracking corrections. The sluggish pitch overdriving type is also discontinuous
with input pulses to stop the unpredictable excess in response. Although the aircraft is
not under complete control, it is not out of control.
High order PIO is a continuous out of control attitude instability, the amplitude ranging
from small to large and potentially destructive. Because the problem is due to inadequate
pilot-vehicle closed loop gain and phase margins, examination of the pitch attitude
frequency response identifies the cause and the solution.
Figure [76] shows the features which separate low and high-order pitch handling. The
area of interest can be confined to the region of phase lags between 180 and 200
degrees which determines the PIO frequency. This arises from the success of the
synchronous pilot (NOR-64-143) in PIO analysis, assuming that any pre-PIO equalization
is abandoned for a pure gain behavior in the undamped or divergent oscillation. The
correct frequency is adopted instantaneously with the stick in phase with the pitch attitude
error and 180 degrees out of phase with the attitude. The stick is not always moved so
purely in practice, but very often the pilot can be seen to apply the stick a little too quickly
and then hold it while waiting for the pitch rate reversal before also reversing the stick.
The tendency of a configuration to PlO can therefore be assessed without using a pilot
model by empirically establishing the range of characteristics found in actual PIO
examples. Enough have now been published to do this with considerable accuracy. An
important feature at the PIO frequency is the response gain. If this is small enough,
dangerous oscillation amplitudes cannot occur, and PIO has not been found where this is
less than 0.1 degrees per pound of stick force. This is not a completely necessary
condition but it is a highly desirable design aim.
PIO's have occurred most frequently, though not exclusively, in the landing flare. The
connection with the commonplace stick pumping is well established. This subconscious
excitation of pitch acceleration in the flare occurs near the same frequency as a PIO. If
the attitude in the oscillation suddenly intrudes into the pilot's awareness, a ready-made
PIO is already in existence. The lower the frequency, the larger is the attitude oscillation
at the usual acceleration amplitude of about 6 deg/sec2, and the more likely the
conversion becomes. This indicates strongly the desirability of a high crossover
frequency through the PIO region.
While an oscillation amplitude of less than 0.5-degrees in the flare will not usually be
noticed, the one significantly more than a degree is very likely to, this or the
corresponding pumping/PIO frequency is not an ideal parameter for correlation. The most
successful has proved to be the rate at which the pitch attitude phase lag increases with
frequency in the PIO lag crossover region, equally applicable to the landing or to target
tracking tasks. By the nature of the attitude frequency response, if the crossover
frequency is low and the attitude attenuates only slowly towards the crossover region, the
phase rate is large. If the frequency is high and there is substantial attenuation, the phase
rate is low. The gain margin is increased, the stick pumping amplitude is reduced, and
the tendency to PIO is decreased automatically by designing a low phase rate into the
control laws.
This simple attitude parameter alone is almost sufficient to quantify the tendency to high
order PIO, and it correlates well with available examples of high order PIO. Figure [77]
shows the trends, with an accuracy good enough to allow Level 1, 2, and 3 boundaries to
be drawn, if desired. For the control law designer ft is enough to aim for a phase rate of
less than 100 degrees per cps and attitude response phase rate of less than 100 degrees
per cps and attitude response smaller than 0.1 deg/lb at the crossover. These
Supersedes pages 267-274 of
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270
MIL-STD-1797A
APPENDIX A
characteristics are a natural feature of low order aircraft whose attitude phase lag
exceeds 180 degrees due to the power control and so could in principle suffer from PIO,
yet do not. Early examples of bobweight PlO were high-order in kind and are found to
have had very large phase rates with the stick free.
For most combat aircraft configurations, consideration of normal acceleration effects
does not improve the PIO analysis. The g at the cockpit is usually attenuated and phase
advanced relative to the cg and will often not reach the 180 degrees lag necessary for
piloted instability. Human sensing of the g response is poor and at the initiation of the PIO
the g may be undetectable. In large aircraft with the cockpit far ahead of the cg, the
heave can have a significant effect and has to be taken into account in the dominant
requirement to optimize the pitch attitude behavior.
Although the attitude to stick force response gain is significant in PIO, there is little
evidence that a damper modifies the pilot's stick phasing in a PIO and only the stiffness
component should be used. Where PIO tendencies exist, they will be exacerbated by a
high stick stiffness. Gradients of 5 to 8 lb/in with forces of 2 to 2.5 lb g have proved to be
extremely satisfactory for [fly-by-wire] aircraft. Designed to the phase rate and gain
margin criteria discussed above, the attitude gain phase rate and gain at the PIO
frequency is only some 0.5 deg/in. In AFFDL-TR-74-9, case 4D had high phase rate and
low PIO gain margin. With a gradient of 22 lb/in and 6.7 lb/g it had an attitude gain of 7
deg/in at the PIO frequency. Not surprisingly it suffered from continuous pitch oscillations
and severe tracking PIO, earning ratings of 9 and 10.
The boundaries in the frequency response criteria of figure [75] are based directly on
these considerations and will eliminate high order PIO. Low order PIO will also be
eliminated by the optimization criteria given above.
REQUIREMENT LESSONS LEARNED
The Smith-Geddes criteria has been used by AFFTC with considerable success for several years. The
criteria has been used to analyze PlOs in the Space Shuttle, the F-15 with CAS-off, the AFTI/F-16, the
AFTI/F-111, the F-15 S/MTD, the YF-22, and the C-17. Application of the criteria to the Space Shuttle
was documented in "Prediction and Occurrence of Pilot-Induced Oscillations in a Flight Test Aircraft” by
Twisdale and Kirsten. In an analysis of three PlOs in the Space Shuttle, the Smith-Geddes criteria
correctly predicted the PIO tendency and closely predicted the frequency of the PIO. For a PIO in landing
flare, the criteria predicted a frequency of 3.5 rad/sec and the frequency of the observed PIO was 3.6
rad/sec. In another PIO at an altitude of 18,000 ft and a speed of 610 ft/sec, the criteria predicted a
frequency of 3.3 rad/sec and the observed frequency was 3.1 rad/sec. In the final example, at a similar
flight condition, a PIO occurred while tracking a cockpit display. When the display dynamics were added
to the analysis, the criteria accurately predicted the observed PIO frequency of 2.0 rad/sec.
A valuable lesson learned in the determination of the criterion frequency, ωc, is found in a Northrop white
paper, "Evaluation of B-2 Susceptibility to Pilot-induced Oscillations" by Margo L. Givens and Frank L.
George, presented at the Flying Qualities Working Group at the 1994 AIAA Atmospheric Flight Mechanics
Conference.
For the most part, the approach taken was as recommended by Ralph Smith in [AFFDLTR-78-154] which presented a straight forward process of evaluation. Exceptions were
made for criterion frequency selection. The recommended method [in AFFDL-TR-78-154]
for criterion frequency selection is based on calculating an average slope of the pitch
attitude-to-controller Bode magnitude plot in the range of 2.0 to 6.0 rad/sec and then
applying this value to the [ωc] formula. This frequency range stipulation was often
inappropriate for the B-2 which has higher break frequencies than those systems
described in the [Smith-Geddes] documentation. Because the validity of a [SmithGeddes] analysis is dependent on the correct selection of the criterion frequency, three
other methods of criterion frequency selection were evaluated.
Supersedes pages 267-274 of
MIL-STD-1797A
271
MIL-STD-1797A
APPENDIX A
The first method, which has been used by Ralph Smith in the past, involved selecting
parameters for a pilot model, closing the pitch attitude loop, and obtaining an nz to αgust
response power spectral density (PSD) in search of resonant frequencies which would be
defined as the criterion frequencies The implementation of this method was unproductive
because no resonant frequencies were found in the B-2 nz PSDs.
The second method adapted the recommended ωc derivation formula to use the slope
calculated after he short period break rather than the average slope in the 2.0 to 6.0
rad/sec frequency range. This method worked quite well for most of the cases. Because
of the occasional case which produced questionable results, a third method of frequency
determination was devised for use as a validity check on criterion frequencies derived
using method 2.
The third approach used typical B-2 pilot pitch stick input frequencies as the criterion
frequencies. These frequencies were determined by calculating PSDs from stick, nz, and
θ time histories of landings and refuelings extracted from flight test data. It was found that
the ωc values calculated with the second method were consistent with these flight data
pilot stick input frequency ranges.
A very good summary report on PlOs is given in NOR-64-143. The following paragraphs from that
reference discuss the causes of PlOs:
There are several ways of looking at the causes of a PIO. One is to catalog all the PIO
situations ever recorded, including all the necessary subsystem details, etc., and then to
say that each combination of vehicle and subsystem when combined with the pilot was
the cause of a PIO. Another way is to note that certain system phenomena such as stickforce-to-control-deflection hysteresis often lead to PIO when other conditions are right. A
third way, and one which seems to transcend the difficulties of the previous two, is to say
that certain inherent human physical limitations are the basic cause for any PIO. This is
not to degrade the human pilot's role but, instead, to emphasize it, because it is unlikely
that any black-box could be devised which is as clever and effective in coping with
unmanageable controlled elements as a skilled pilot. Were it not for the pilot's versatile
gain adaptability, many flight conditions would be unstable. But there is a limit to the
rapidity with which the human can adapt, and this can sometimes lead to a PIO.
When referred to the pilot, then, the basic causes of PIO seem to fall into the following
categories:
1. Incomplete pilot equalization
a. Incomplete training
b. Inappropriate transfer of adaptation (i.e., carry over of improper techniques from
another aircraft)
2. Excessive-demands on pilot adaptation
a. Required gain, lead, or lag lie outside the range of normal capabilities
b. Rate of adaptation is too slow to preclude oscillation
c.
Inadequate capability to cope with system nonlinearities
3. Limb-manipulator coupling
a. Impedance of neuromuscular system (including limb) on control stick or pedals
changes feel system dynamics
Supersedes pages 267-274 of
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272
MIL-STD-1797A
APPENDIX A
b. Motion-induced limb force feedback (e.g., arm becomes a bobweight)
Table XIV, from NOR-64-143, lists some known PIO cases and their probable causes for then-current
(early 1960s) aircraft. The causes are equally relevant for modern aircraft, and the lessons learned from
the cases listed are valuable in preventing PlOs.
5.2.2 Pilot-in-the-loop oscillations - verification.
Verification shall be by analysis, simulation, and flight test.
VERIFICATION RATIONALE (5.2.2)
It would be an easy matter for the engineers of the procuring agency to ascertain compliance with this
paragraph without relying on pilot/vehicle analysis methods. For example, ωc and the specified phase lag
can easily be obtained from simulator or in-flight time histories (ground-based simulations will not show up
acceleration-dependent PIO tendencies). Nonetheless, analytical estimates can - and should - be made
by the airframe manufacturer as part of the design evolution. For flight evaluation, the PIO tendency
classification scale of figure 12 will be helpful.
VERIFICATION GUIDANCE
The user should refer to AFFDL-TR-77-57, AFFDL-TR-78-154, and AFWAL-TR-81-3090 when applying
the quantitative requirement. PlOs are associated with abrupt maneuvers and precise tracking as in airto-air gunnery, formation flying, flare, and touchdown. PlOs observed in flight are often not obtained in
ground-based simulators, even ones with some motion. Tight, aggressive pilot control action will tend to
bring on any PIO tendencies. High sensitivity to control inputs is often a factor. Some pilots are more PIOprone than others, depending upon piloting technique.
VERIFICATION LESSONS LEARNED
These requirements are an attempt to catch and correct any PIO tendencies as early as possible in the
design, when changes are easiest and least costly to make. They also have been found helpful in
identifying PIO tendencies in flight and determining fixes.
Supersedes pages 267-274 of
MIL-STD-1797A
273/274
MIL-STD-1797A
APPENDIX A
TABLE XIV. Classification of some known PIO cases (from NOR-64-143).
Examples shown as: SPECIES (Aircraft); Critical Subsystem; Critical Flight Condition: Remarks
TYPE
CLASS
I. LINEAR
II. SERIES NONLINEAR ELEMENTS
III. SUBSIDIARY FEEDBACK NONLINEAR
ELEMENTS
IMPROPER SIMULATION; D, V; a:
Abnormally high value of 1/Tθ2 and low ζωsp
led to zero ζsp when regulating large
disturbances.
PORPOISING (SB2C-1); F; c: Hysteresis in
stick versus elevator deflection resulted in lowfrequency speed and climb oscillations.
BOBWEIGHT BREAKOUT (A4D-1, T-38A):
GCA-INDUCED PHUGOID (C-97); D; c, b:
Lag from radar-detected error to voice
command led to unstable closed-loop
phugoid mode.
J. C. MANEUVER (F-86D, F-100C); F, S; a:
Valve friction plus compliant cabling resulted in
large oscillations at short period.
LOSS OF PITCH DAMPER
ARM ON STICK (A4D-1, T-38A); F; a: Arm
mass increases feel system inertia, leads via
B feedback to unstable coupling with shortperiod dynamics if pilot merely hangs loosely
onto stick after a large input.
PITCH-UP (XF-104, F-101B, F-102A); V; c:
Unstable kink in M(α) curve led to moderateperiod oscillations of varying amplitudes
(depending on extent and nature of the kink)
during maneuvers near the critical angle-ofattack.
F, B; a: At high-g maneuvers the bobweight
overcomes system friction and reduces
apparent damping of the aircraft in response
to force inputs, resulting in large oscillations at
short-period.
PITCH
LANDING PIO (X-15); S; b: Closed loop around
elevator rate-limiting caused moderate
oscillations at short period.
ωφ/ωd EFFECT (X-15, NT-33A, F-101B,
LOSS OF YAW DAMPER
F-106A, KC-135A, B-58); V; c: Zeros of
roll/aileron transfer function are higher than
Dutch roll frequency,
LATERALDIRECTIONAL
ω φ / ω d > 1.0 ,
leading to closed-loop instability at low ζ d
conditions.
BORESIGHT OSCILLATIONS (F-5A);
D, V; c: Spiral roll mode driven unstable if roll
information is degraded during gunnery.
YAW
FUEL SLOSH SNAKING (KC-135A, T-37A);
V; c: Fuel slosh mode couples with Dutch roll
mode when rudder used to stop yaw
oscillation.
TRANSONIC SNEAKING (A3D); V, F; a, c:
Separation over rudder causes control reversal
for small deflections, leading to limit cycle if
rudder used to damp yaw oscillations.
ROLL
NONE KNOWN
PILOT-INDUCED CHATTER (F-104B); A; c:
Small limit cycle due to damper aggravated
whenever pilot attempted to control it.
* Critical Subsystems:
D = Display
F = Feel system (except B)
B = Bobweight
** Critical Flight Conditions:
a = Low altitude, near-sonic Mach
b = Landing approach and takeoff
c = Cruise
S = Power servo actuator
V = Vehicle (airframe)
A = Augmentor (damper)
275
MIL-STD-1797A
APPENDIX A
4.2.3 Residual pitch oscillations.
In calm air, any sustained residual oscillations shall not interfere with the pilot's ability to perform the tasks
required in service use of the aircraft. For Levels 1 and 2, oscillations in normal acceleration at the pilot
station greater than will be considered excessive for any Flight Phase. These requirements shall apply
with the pitch control fixed and with it free.
REQUIREMENT RATIONALE (4.2.3)
The requirement prohibits limit cycles in the control system or structural oscillations that might
compromise tactical effectiveness, cause pilot discomfort, etc. This requirement may be considered a
relaxation of the requirement in 4.2.1 for positive damping at all magnitudes of oscillation. Its intent is to
recognize thresholds below which damping is immaterial.
REQUIREMENT GUIDANCE
The related MIL-F-4785C requirement is paragraph 3.2.2.1.3.
The recommended value is 0.02g. Given the proper data, this threshold could be made a function of
frequency in order to correspond more closely with human perception.
REQUIREMENT LESSONS LEARNED
Allowable normal acceleration oscillations have been decreased to 0.02g from the 0.05g of MIL-F-8785C.
This is based on flight test experience with the B-1 (AFFTC-TR-79-2), which encountered limit cycle
oscillations during aerial refueling, subsonic and supersonic cruise. A primary contributor was identified to
be mechanical hysteresis in the pitch system. According to AFFTC-TR-79-2, "Flying qualities were initially
undesirable due to this limit cycle." Normal acceleration transients in cruise were about 0.05-0.12 g, as
figure 94 shows. The limit cycle was eliminated by installation of a mechanical shaker (dither) vibrating at
20 Hz.
5.2.3 Residual pitch oscillations - verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.3)
Limit cycle amplitude depends on characteristics of the actual hardware and software, and so may be
different in simulations than in actual flight. Measurements of normal acceleration at the pilot's station
should be made in the course of test flight to meet the other flying quality requirements.
VERIFICATION GUIDANCE
Residual oscillations are limit cycles resulting from nonlinearities such as friction and poor resolution.
Negative static stability will contribute and low damping may augment the amplitude. Thus high speed,
high dynamic pressure or high attitude may be critical. Residual oscillations are most bothersome in
precision tasks.
VERIFICATION LESSONS LEARNED
REPRINTED WITHOUT CHANGE
276
MIL-STD-1797A
APPENDIX A
4.2.5 Pitch trim changes.
The pitch trim changes caused by operation of other control devices shall not be so large that a peak
pitch control force in excess of 10 pounds for center-stick controllers or 20 pounds for wheel controllers is
required when such configuration changes are made in flight under conditions representative of
operational procedure. Generally, the conditions of table IV will suffice for determination of compliance
with this requirement. With the aircraft trimmed for each specified initial condition, and no retrimming, the
peak force required to maintain the specified parameter constant following the specified configuration
change shall not exceed the stated value for a time interval of at least 5 seconds following the completion
of the pilot action initialing the configuration change. The magnitude and rate of trim change subsequent
to this time period shall be easily trimmable by use of the normal trimming devices. These requirements
define Level 1. For Levels 2 and 3, the allowable forces are increased by 50 percent.____________
REQUIREMENT RATIONALE (4.2.5)
These frequently encountered pitch trim changes, if too large, can add to pilot workload at critical times
during a mission.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.6.3.1.
Table XV gives the recommended conditions (For aircraft with variable-sweep wings, additional
requirements should be imposed consistent with operational employment of the vehicle. Thrust reversing
and other special features also need to be considered). These are the trim changes that, when larger
than the limits specified, have been bothersome in the past. Crossfeeds and feedbacks in the stability and
control augmentation system generally will reduce the magnitude of these trim changes. Wing downwash
and vertical placement of the engines are two of the determining factors. For thrust reversing,
configuration-dependent aerodynamics play an important role.
4.1.13 gives additional general trim requirements.
REQUIREMENT LESSONS LEARNED
The direction of the trim change can also be important, producing either helpful or unfavorable coupling.
In any case the magnitude should not be excessive.
5.2.5 Pitch trim changes - verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.5)
The evaluation should be made in the manner expected in operational practice, rather than necessarily
holding everything else constant.
VERIFICATION GUIDANCE
Initial trim conditions are listed in table XV.
VERIFICATION LESSONS LEARNED
Supersedes pages 281 – 282 of
MIL-STD-1797A
281
MIL-STD-1797A
APPENDIX A
TABLE XV. Pitch trim change conditions.
Initial Trim Conditions
Flight
Phase
Attitude
Speed
Landing
Gear
High-lift
Devices
& Wing
Flaps
Thrust
Configuration
Change
Parameter to
be
held
constant
Approach
ho min
Normal
pattern
entry speed
Up
Up
TLF
Gear down
Altitude and
airspeed*
2
Up
Up
TLF
Gear down
Altitude
3
Down
Up
TLF
Extend highlift devices and
wing flaps
Altitude and
airspeed*
4
Down
Up
TLF
Extend highlift devices and
wing flaps
Altitude
5
Down
Down
TLF
Idle thrust
Airspeed
Down
Down
TLF
Extend
approach drag
device
Airspeed
Down
Down
TLF
Takeoff thrust
Airspeed
Down
Down
TLF
Takeoff thrust
plus
normal
cleanup
for
wave-off (goaround)
Airspeed
Down
Take-off
Take-off
thrust
Gear up
Pitch attitude
Minimum
flap-retract
speed
Up
Take-off
Take-off
thrust
Retract highlift devices and
wing flaps
Airspeed
Speed for
level flight
Up
Up
MRT
Idle thrust
Pitch attitude
12
Up
Up
MRT
Actuate
celeration
devices
13
Up
Up
MRT
Maximum
augmented
thrust
Up
Up
TLF
Actuate
deceleration
device
1
6
Vomin
7
8
Approach
9
Takeoff
Vomin
10
11
Cruise
and airto-air
combat
14
homin and
homax
Speed for
best range
de-
*Throttle setting may be changed during the maneuver.
Notes: - Auxiliary drag devices are initially retracted, and all details of configuration not specifically mentioned
are normal for the Flight Phase.
Supersedes pages 281 – 282 of
MIL-STD-1797A
282
MIL-STD-1797A
APPENDIX A
- If power reduction is permitted in meeting the deceleration requirements established for the mission,
actuation of the deceleration device in #12 and #14 shall be accompanied by the allowable power
reduction.
Supersedes pages 281 – 282 of
MIL-STD-1797A
283
MIL-STD-1797A
APPENDIX A
5.5.1.5 Time delay - verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.5.1.5)
In the end, flight test data or a flight-verified analytical model should be used to verify compliance.
A control surface rate limit may increase the equivalent time delay or roll-mode time constant as a
function of the size of command.
VERIFICATION GUIDANCE
Appropriate values of τ ep will require equivalent system matching, as discussed above. See guidance for
4.5.1.1.
VERIFICATION LESSONS LEARNED
REPRINTED WITHOUT CHANGE
421
MIL-STD-1797A
APPENDIX A
4.5.2 Pilot-in-the-Ioop roll oscillations.
There shall be no tendency for sustained or uncontrollable roll oscillations resulting from efforts of the
pilot to control the aircraft. The phase angle of the bank angle frequency response to roll stick force at the
criterion frequency, w shall be greater than or equal to ________. Furthermore, the requirements of
4.5.1.1, 4.5.1.3, 4.5.1.4, 4.5.1.5, 4.5.8.1, 4.5.9.2, and 4.5.9.3 must be met.
REQUIREMENT RATIONALE (4.5.2)
This roll-axis requirement is stated in addition to the general requirement of 4.1.11.6 to emphasize its
importance for the roll axis and to allow incorporation of a more quantitative requirement.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.3.3.
Recommended values:
The recommended minimum phase angle of the bank angle frequency response to the roll stick force at
the criterion frequency, ωc, is -180°.
The value of ωc is determined as follows:
ωc = (0.24 rad-oct/dB-sec) S + 6.0 rad/sec
where S is the average slope of
φ(s )
in dB/oct over the interval from 1 to 6 rad/sec.
Fas (s )
This requirement is the Smith-Geddes PIO criteria extended to the lateral-directional axis. The origins of
the Smith-Geddes criteria are traced in 4.2.2 Requirement Guidance, and discussions of the calculation
of S and ωc can be found there. Application in the lateral-directional axis is similar to that in the
longitudinal axis, except that it is applied to φ/Fas instead of to θ/Fes.
The statement that requirements 4.5.1.1, 4.5.1.3, 4.5.1.4, 4.5.1.5, 4.5.8.1, 4.5.9.2, and 4.5.9.3 must also
be met would seem to be redundant, since these are already requirements. However, recent history
would seem to indicate that, because the term PIO does not appear in these requirements, the
importance of these requirements; in precluding PIO is not appreciated. Many recent PIO incidents can
be traced directly to problems addressed by these requirements. Therefore, these requirements are
repeated here in the PIO requirement to insure that their significance in precluding PIO tendencies is
understood.
REQUIREMENT LESSONS LEARNED
The extension of the Smith-Geddes criteria to the lateral-directional axis was developed in AFWAL-TR81-3090. In AFWAL-TR-81-3090, the lateral-directional criteria was used to analyze the YF-16, the X-15,
and the M2-F2 and M2-F3 lifting bodies and also compared with results from handling qualities research
projects with variable-stability aircraft: an approach and landing evaluation with the NT-33 (AFWAL-TR81-3116), an investigation of reentry vehicle lateral-directional dynamics on the NT-33 (WADD-TR-61147), and one configuration from lateral-directional studies on the Princeton Navion (Princeton University
Report No. 727). Most of the data support the PIO criteria, and, in those cases where PIO was predicted
but not encountered, handling qualities were usually poor.
See 4.2.2 for discussion of applicable considerations and data, in that case directed at longitudinal PlOs
in general. The M2-F2 lifting body (NASA-TN-D-6496) encountered several divergent PlOs during flight
testing. The primary cause was found to be the coupled roll subsidence/spiral mode (see Lessons
Learned for 4.5.1.3).
Supersedes page 422 of MIL-STD-1797A
422
MIL-STD-1797A
APPENDIX A
Another cause of observed lateral PIO tendencies is the ω φ / ω d effect noted and explained in figure 158
and also in Norair Rpt No. NOR-64-143. Another prevalent cause is associated with control-surface rate
saturation. In this case the pilot tries to apply lateral control at a rate greater than the maximum surface
rate, thereby getting out of phase if tight tracking is attempted. The quantitative aspects of such ratelimiting are given in the appendix of Norair Rpt No. NOR-64-143 and involve gain and phase decrements
that are functions of the ratio of commanded to saturation rate.
PlOs on recent aircraft have been related to roll responses which are both too low (F-18) and too high
(YF-16). These cases are discussed under 4.5.8.1 and 4.5.9.3. Control sensitivity, control surface rate
limiting, control surface saturation, and (equivalent) time delay are critical factors in roll PIO.
5.5.2 Pilot-in-the-loop roll oscillations - verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.5.2)
This requirement should apply to all flight conditions and tasks, and to all Levels, since zero or negative
closed-loop damping is to be avoided under all flight conditions and failure states.
VERIFICATION GUIDANCE
The existence of a PIO tendency is difficult to assess. A high-stress task such as approach and landing
with a lateral offset, air-to-air tracking, or terrain following, may reveal PIO proneness. Demanding
tracking tasks, aggressive control, sensitive response, proverse yaw, low dutch roll damping and long
equivalent time delay are factors varying with flight condition which may tend to incite roll PlOs. Lateral
acceleration induced on the pilot in rolling may contribute.
VERIFICATION LESSONS LEARNED
In a number of cases optimization of p/Fas in a fixed-base simulator has resulted in gross oversensitivity in
actual flight.
Supersedes page 422 of MIL-STD-1797A
422a
MIL-STD-1797A
APPENDIX A
2. Multiply the result by N′δrp L ′δas , i.e., δ′rp (3) = YCF(3) ⋅ N′δrp L ′δas
3. Compare δ′rp (3) with table XLIV
7. If 0.03 < N′δas L ′δas ≤ 0.07, utilize the more conservative result from steps 5 and 6.
8.
If the configuration does not meet the requirements, see figure 249 and table XLVII to
determine the type of expected piloting problems.
9. In the end, the transfer functions should be identified from flight data.
VERIFICATION GUIDANCE
The flight testing to obtain ∆β and φt command should cover the range of operational altitudes and service
speeds. As with roll rate oscillations (4.5.1.4), the critical flight conditions for compliance with this
requirement should in general become apparent during the roll performance testing of 4.5.5.1. The most
important flight conditions for compliance demonstration of either alternative are those with low |φ/β|d less
than 6.
An approximation for |φ/β|d is
φ
β
(
d
)
(
)
 L ′β + Yβ L ′r 2 + 2ζ d ω dL ′r L ′β + Yβ L ′r + ω 2d L ′r 2 

≈
2
2
2


′
′
L
2
L
ω
+
ζ
ω
+
ω
d
p
d
d
p
d


[
]
12
≈
12
2


2


 C′l 
′
ρgb
1 − ρgb C + 1 C′ − k z Cl β  C − 1 C′  C′l r +
 r  C′ 
y
n
L
n
nβ 
4( W / S)  β 2k 2z r k 2x C′nβ  1 2k 2z p  C′l β 8( W / S)k 2z  C′l β 
k 2zC′l β 







2 
2 C′ 
2 C′2
 C′l
k 2xC′nβ 

l
l
ρ
ρ
gb
k
1
k
1
gb
k
β
p
z 
z
z
 1−

 CL −
 p +
′
′
+
−
C
C
C
y
n
n

4( W / S) k 2x  β 2k 2z r k 2x C′nβ  1 2k 2z p  C′nβ 8( W / S) k 2x C′nβ 




VERIFICATION LESSONS LEARNED
REPRINTED WITHOUT CHANGE
569
MIL-STD-1797A
APPENDIX A
4.6.3 Pilot-in-the-loop yaw oscillations.
There shall be no tendency for sustained or uncontrollable yaw oscillations resulting from efforts of the
pilot to control the aircraft in the air or on the ground. Furthermore, the requirements of 4.6.2 must be met.
REQUIREMENT RATIONALE (4.6.3)
This requirement, in addition to the general requirement of 4.1.11.6, is inserted to provide more specific
criteria for any task that might involve high-bandwidth control in yaw or lateral acceleration. An example
might be yaw pointing for fine tracking.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.3.3.
The statement that requirement 4.6.2 must also be met would seem to be redundant, since this is already
a requirement. However, as with the corresponding requirements in the pitch and roll axes, because the
term PIO does not appear in requirement 4.6.2, the importance of this requirement in precluding PIO may
not be appreciated. Therefore, requirement 4.6.2 is repeated here in the PIO requirement to insure that its
significance in precluding PIO tendencies is understood.
Due to the lack of a reliable quantitative measure, the requirement is written in terms of subjective
evaluations. It is of course hoped that meeting the (other) quantitative requirements of this standard will
prevent a lateral PIO. This requirement is identical to the roll-axis requirement of 4.5.2.
This requirement should apply to all flight conditions and tasks, and to all Levels, since zero or negative
closed-loop damping is to be avoided under any flight condition or failure state. High-bandwidth yawcontrol tasks are uncommon. The dynamic yaw response requirement (4.6.2.1) is designed to account for
the need of rudder pedal in rolling, but may not coverall contingencies. Some direct sideforce modes may
involve high-bandwidth yaw control; see AFWAL-TR-81-3027 and Sammonds, et al., for example.
REQUIREMENT LESSONS LEARNED
The pitch-axis PIO requirement, 4.2.2, discusses some causes of PlOs. Factors known to contribute to
lateral-directional PlOs are large effective or equivalent time delays, excessive friction or hysteresis in the
flight control system, and the “ ω φ / ω d ” effect described at length in AFFDL-TR-69-72 (See 4.5.1.4
discussion). Depending upon the cause, ground-based simulation may or may not prove a useful
investigation technique - often it does not.
5.6.3 Pilot-in-the-loop yaw oscillations - verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.6.3)
A precision closed-loop task, performed aggressively, is needed.
VERIFICATION GUIDANCE
The existence of a PIO tendency is difficult to assess. Therefore, no specific flight conditions or tasks are
recommended, though a high-stress task such as approach and landing with a lateral offset, terrain
following, air-to-ground tracking, or in-flight refueling (receiver) may reveal PIO proneness.
VERIFICATION LESSONS LEARNED
Supersedes page 570 of MIL-STD-1797A
570
INCH-POUND
MIL-STD-1797A
30 January 1990
SUPERSEDING
MIL-STD-1797(USAF)
31 March 1987
MILITARY STANDARD
FLYING QUALITIES
OF
PILOTED AIRCRAFT
AMSC N/A
FSC 15GP
DISTRIBUTION STATEMENT D. Distribution authorized to the Department of Defense and DoD
contractors only; contains critical technology; (3 February 1986). Other requests shall be referred to
ASD/ENES, Wright-Patterson AFB OH 45433-6503.
MIL-STD-1797A
DEPARTMENT OF DEFENSE
WASHNGTON DC 20402
Flying Qualities of Piloted Aircraft
1. This military standard is approved for use by all Departments and Agencies of the Department of
Defense.
2. Beneficial comments (recommendations, additions, deletions) and any pertinent data which may be of
use in improving this document should be addressed to: ASD/ENES, Wright-Patterson AFB, OH 454336503, by using the self-addressed Standardization Document Improvement Proposal (DD Form 1426)
appearing at the end of this document or by letter.
WARNING
INFORMATION SUBJECT TO EXPORT CONTROL LAWS
This document contains information subject to the International Traffic in Arms Regulation (ITAR) and/or
the Export Administration Regulation (EAR) of 1979 which may not be exported, released, or disclosed to
foreign nationals inside or outside the United States without first obtaining an export license. A violation of
the ITAR or EAR may be subject to penalty of up to 10 years imprisonment and a fine of $100,000 under
22 U.S.C. or Section 2410 of the Export Administration Act of 1979. Include this notice with any
reproduced portion of this document.
DESTRUCTION NOTICE. For classified documents, follow the procedures in DoD 5220.22-M Industrial
Security Manual, Section 11-19 or DoD 5200.1-R, Information Security Program Regulation, Chapter IX.
For unclassified, limited documents, destroy by any method that will prevent disclosure of contents or
reconstruction of the document.
ii
MIL-STD-1797A
FOREWORD
This standard is intended for use with fixed-wing aircraft supported primarily by aerodynamic force rather
than engine thrust. It also covers the handling characteristics of aircraft under piloted control on the
ground, and may be used with powered-lift aircraft in aerodynamic flight (above the conversion speed,
Vcon). This standard also applies to piloted transatmospheric flight when flight depends upon aerodynamic
lift and/or air breathing propulsion systems. Flying qualities of military rotorcraft are specified in MIL-H8501, while flying qualities in V/STOL flight are the subject of MIL-F-83300.
For further background information, see Appendix C.
iii
MIL-STD-1797A
CONTENTS
Paragraph
Page
1.
1.1
1.2
SCOPE ...................................................................................................................................... 1
Purpose. .................................................................................................................................... 1
Applicability................................................................................................................................ 1
2.
2.1
2.1.1
2.1.2
2.2
2.3
APPLICABLE DOCUMENTS ................................................................................................... 1
Government documents............................................................................................................ 1
Specifications, standards, and handbooks................................................................................ 1
Other Government documents, drawings, and publications. .................................................... 1
Non-Government publications................................................................................................... 1
Order of precedence. ................................................................................................................ 1
3.
3.1
3.2
3.3
3.4
3.4.1
3.4.2
3.4.3
3.4.4
3.4.5
3.4.6
3.4.7
3.5
DEFINITIONS............................................................................................................................ 2
Aircraft classification and operational missions. ....................................................................... 2
Flight Phase Categories............................................................................................................ 2
Levels and qualitative suitability of flying qualities. ................................................................... 2
Parameters................................................................................................................................ 2
General terms ........................................................................................................................... 2
Speeds ...................................................................................................................................... 3
Thrust and power ...................................................................................................................... 5
Control parameters ................................................................................................................... 6
Longitudinal parameters............................................................................................................ 6
Lateral-directional parameters .................................................................................................. 8
Atmospheric disturbance parameters ..................................................................................... 16
Terms used in high angle of attack requirements ................................................................... 18
4.
4.1
4.1.1
4.1.2
4.1.3
4.1.4
4.1.4.1
4.1.4.2
4.1.4.3
4.1.5
4.1.6
4.1.6.1
4.1.6.2
4.1.6.3
4.1.7
4.1.7.1
4.1.7.2
4.1.7.3
4.1.7.4
4.1.7.5
4.1.7.6
4.1.8
4.1.8.1
4.1.8.2
4.1.9
4.1.10
4.1.11
4.1.11.1
REQUIREMENTS ................................................................................................................... 19
General requirements ............................................................................................................. 19
Loadings.................................................................................................................................. 19
Moments and products of inertia............................................................................................. 19
Internal and external stores..................................................................................................... 19
Flight Envelopes...................................................................................................................... 19
Operational Flight Envelopes. ................................................................................................. 19
Service Flight Envelopes. ........................................................................................................ 19
Permissible Flight Envelopes. ................................................................................................. 19
Configurations and States of the aircraft................................................................................. 19
Aircraft Normal States. ............................................................................................................ 21
Allowable Levels for Aircraft Normal States. ........................................................................... 21
Flight outside the Service Flight Envelopes. ........................................................................... 21
Ground operation..................................................................................................................... 21
Aircraft Failure States.............................................................................................................. 21
Allowable Levels for Aircraft Failure States............................................................................. 21
Aircraft Special Failure States. ................................................................................................ 21
Probability calculation.............................................................................................................. 21
Generic failure analysis. .......................................................................................................... 23
When Levels are not specified. ............................................................................................... 23
Failures outside the Service Flight Envelopes. ....................................................................... 23
Dangerous flight conditions..................................................................................................... 23
Warning and indication. ........................................................................................................... 23
Devices for indication, warning, prevention, and recovery...................................................... 24
Interpretation of subjective requirements................................................................................ 24
Interpretation of quantitative requirements.............................................................................. 24
General flying qualities requirements ...................................................................................... 24
Buffet. ...................................................................................................................................... 24
iv
MIL-STD-1797A
CONTENTS
Paragraph
4.1.11.2
4.1.11.3
4.1.11.4
4.1.11.5
4.1.11.6
4.1.11.7
4.1.11.8
4.1.12
4.1.12.1
4.1.12.2
4.1.12.3
4.1.12.4
4.1.12.5
4.1.12.6
4.1.12.7
4.1.12.8
4.1.12.9
4.1.12.10
4.1.12.11
4.1.13
4.1.13.1
4.1.13.2
4.1.13.3
4.1.13.4
4.1.13.5
4.1.13.6
4.2
4.2.1
4.2.1.1
4.2.1.2
4.2.2
4.2.3
4.2.4
4.2.5
4.2.6
4.2.6.1
4.2.6.2
4.2.7
4.2.7.1
4.2.7.2
4.2.7.2.1
4.2.7.3
4.2.7.4
4.2.8
4.2.8.1
4.2.8.2
4.2.8.3
4.2.8.4
4.2.8.5
4.2.8.6
4.2.8.6.1
4.2.8.6.2
4.2.8.6.3
Page
Release of stores..................................................................................................................... 24
Effects of armament delivery and special equipment.............................................................. 24
Failures. ................................................................................................................................... 24
Control margin. ........................................................................................................................ 25
Pilot-induced oscillations (PIO)................................................................................................ 25
Residual oscillations. ............................................................................................................... 25
Control cross-coupling............................................................................................................. 25
General flight control system characteristics........................................................................... 25
Control centering and breakout forces. ................................................................................... 25
Cockpit control free play. ......................................................................................................... 26
Adjustable controls. ................................................................................................................. 26
Rate of control displacement. .................................................................................................. 26
Dynamic characteristics........................................................................................................... 26
Damping. ................................................................................................................................. 26
Transfer to alternate control modes. ....................................................................................... 26
Flight control system failures. .................................................................................................. 26
Augmentation systems. ........................................................................................................... 26
Auxiliary dive recovery devices. .............................................................................................. 27
Direct force controllers............................................................................................................. 27
General trim requirements....................................................................................................... 27
Trim system irreversibility. ....................................................................................................... 27
Rate of trim operation. ............................................................................................................. 27
Stalling of trim systems............................................................................................................ 27
Transients and trim changes. .................................................................................................. 27
Trim for asymmetric thrust....................................................................................................... 28
Automatic trim system. ............................................................................................................ 28
Flying qualities requirements for the pitch axis. ...................................................................... 28
Pitch attitude dynamic response to pitch controller................................................................. 28
Long-term pitch response........................................................................................................ 28
Short-term pitch response. ...................................................................................................... 28
Pilot-induced pitch oscillations. ............................................................................................... 28
Residual pitch oscillations. ...................................................................................................... 28
Normal acceleration at pilot station......................................................................................... 28
Pitch trim changes................................................................................................................... 28
Pitch axis response to other inputs ......................................................................................... 29
Pitch axis response to failures, controls free........................................................................... 29
Pitch axis response to configuration or control mode change................................................. 29
Pitch axis control power .......................................................................................................... 31
Pitch axis control power in unaccelerated flight. ..................................................................... 31
Pitch axis control power in maneuvering flight. ....................................................................... 31
Load factor response............................................................................................................... 31
Pitch axis control power in takeoff........................................................................................... 31
Pitch axis control power in landing. ......................................................................................... 31
Pitch axis control forces .......................................................................................................... 31
Pitch axis control forces--steady-state control force per g. ..................................................... 31
Pitch axis control forces--transient control force per g. ........................................................... 31
Pitch axis control forces--control force variations during rapid speed changes. ..................... 32
Pitch axis control forces--control force vs. control deflection. ................................................. 32
Pitch axis control breakout forces. .......................................................................................... 32
Pitch axis control force limits ................................................................................................... 32
Pitch axis control force limits--takeoff...................................................................................... 32
Pitch axis control force limits--landing. .................................................................................... 32
Pitch axis control force limits--dives. ....................................................................................... 32
v
MIL-STD-1797A
CONTENTS
Paragraph
4.2.8.6.4
4.2.8.6.5
4.2.8.6.6
4.2.8.7
4.2.9
4.2.9.1
4.2.9.2
4.3
4.3.1
4.3.1.1
4.3.1.2
4.3.2
4.3.3
4.3.3.1
4.3.3.2
4.3.4
4.4
4.4.1
4.4.1.1
4.5
4.5.1
4.5.1.1
4.5.1.2
4.5.1.3
4.5.1.4
4.5.1.5
4.5.2
4.5.3
4.5.4
4.5.5
4.5.6
4.5.7
4.5.7.1
4.5.7.2
4.5.8
4.5.8.1
4.5.8.2
4.5.8.3
4.5.8.4
4.5.8.5
4.5.8.6
4.5.9
4.5.9.1
4.5.9.2
4.5.9.3
4.5.9.4
4.5.9.5
4.5.9.5.1
4.5.9.5.2
4.5.9.5.3
4.5.9.5.4
4.5.9.5.5
4.5.9.5.6
Page
Pitch axis control force limits-sideslips. ................................................................................... 32
Pitch axis control force limits-failures. ..................................................................................... 33
Pitch axis control force limits-control mode change. ............................................................... 33
Pitch axis trim systems. ........................................................................................................... 33
Pitch axis control displacements ............................................................................................. 33
Pitch axis control displacements-takeoff. ................................................................................ 33
Pitch axis control displacements-maneuvering. ...................................................................... 33
Flying qualities requirements for the normal (flight path) axis ................................................ 33
Flight path response to attitude change.................................................................................. 33
Transient flight path response to attitude change. .................................................................. 33
Steady-state flight path response to attitude change. ............................................................. 33
Flight path response to designated flight path controller. ....................................................... 34
Flight path control power......................................................................................................... 34
Control power for designated primary flight path controller..................................................... 34
Control power for designated secondary flight path controller. ............................................... 34
Flight path controller characteristics. ...................................................................................... 34
Flying qualities requirements for the longitudinal (speed) axis ............................................... 34
Speed response to attitude changes. ..................................................................................... 34
Speed response to attitude changes-relaxation in transonic flight.......................................... 34
Flying qualities requirements for the roll axis.......................................................................... 34
Roll response to roll controller................................................................................................. 34
Roll mode................................................................................................................................. 34
Spiral stability........................................................................................................................... 34
Coupled roll-spiral oscillation................................................................................................... 34
Roll oscillations........................................................................................................................ 35
Roll time delay. ........................................................................................................................ 35
Pilot-induced roll oscillations. .................................................................................................. 35
Linearity of roll response to roll controller. .............................................................................. 35
Lateral acceleration at the pilot station. .................................................................................. 35
Roll characteristics in steady sideslip. .................................................................................... 35
Roll axis control for takeoff and landing in crosswinds. .......................................................... 35
Roll axis response to other inputs ........................................................................................... 35
Roll axis response to augmentation failures............................................................................ 35
Roll axis response to configuration or control mode change. ................................................. 35
Roll axis control power ............................................................................................................ 36
Roll axis response to roll control inputs................................................................................... 36
Roll axis control power in steady sideslips. ............................................................................. 36
Roll axis control power in crosswinds...................................................................................... 36
Roll axis control power for asymmetric thrust.......................................................................... 36
Roll axis control power in dives and pullouts........................................................................... 36
Roll axis control power for asymmetric loading. ...................................................................... 36
Roll axis control forces and displacements............................................................................. 37
Roll control displacements....................................................................................................... 37
Roll axis control forces to achieve required roll performance. ................................................ 37
Roll axis control sensitivity....................................................................................................... 37
Roll axis control centering and breakout forces. ..................................................................... 37
Roll axis control force limits ..................................................................................................... 37
Roll axis control force limits in steady turns. ........................................................................... 37
Roll axis control force limits in dives and pullouts. .................................................................. 37
Roll axis control force limits in crosswinds. ............................................................................. 37
Roll axis control force limits in steady sideslips....................................................................... 37
Roll axis control force limits for asymmetric thrust. ................................................................. 37
Roll axis control force limits for failures. .................................................................................. 37
vi
MIL-STD-1797A
CONTENTS
Paragraph
Page
4.5.9.5.7
4.6
4.6.1
4.6.1.1
4.6.1.2
4.6.1.3
4.6.2
4.6.3
4.6.4
4.6.5
4.6.5.1
4.6.5.2
4.6.5.3
4.6.6
4.6.6.1
4.6.6.2
4.6.6.3
4.6.7
4.6.7.1
4.6.7.2
4.6.7.3
4.6.7.4
4.6.7.5
4.6.7.6
4.6.7.7
4.6.7.8
4.6.7.9
4.6.7.10
4.6.7.11
4.7
4.7.1
4.8
4.8.1
4.8.2
4.8.3
4.8.4
4.8.4.1
4.8.4.2
4.8.4.2.1
4.8.4.2.2
4.8.4.2.3
4.8.4.2.4
4.8.4.3
4.8.4.3.1
4.8.4.3.2
4.9
4.9.1
4.9.2
4.9.3
Roll axis control force limits for configuration or control mode change................................... 38
Flying qualities requirements for the yaw axis ........................................................................ 38
Yaw axis response to yaw and side-force controllers ............................................................. 38
Dynamic lateral-directional response. ..................................................................................... 38
Steady sideslips....................................................................................................................... 38
Wings-level turn....................................................................................................................... 38
Yaw axis response to roll controller. ....................................................................................... 39
Pilot-induced yaw oscillations. ................................................................................................ 39
Yaw axis control for takeoff and landing in crosswinds. ......................................................... 39
Yaw axis response to other inputs .......................................................................................... 39
Yaw axis response to asymmetric thrust................................................................................. 39
Yaw axis response to failures.................................................................................................. 40
Yaw axis response to configuration or control mode change.................................................. 40
Yaw axis control power. .......................................................................................................... 40
Yaw axis control power for takeoff, landing, and taxi. ............................................................. 40
Yaw axis control power for asymmetric thrust......................................................................... 41
Yaw axis control power with asymmetric loading. ................................................................... 41
Yaw axis control forces. .......................................................................................................... 41
Yaw axis control force limits in rolling maneuvers................................................................... 41
Yaw axis control force limits in steady turns............................................................................ 41
Yaw axis control force limits during speed changes. .............................................................. 41
Yaw axis control force limits in crosswinds. ............................................................................ 41
Yaw axis control force limits with asymmetric loading. ........................................................... 42
Yaw axis control force limits in dives and pullouts. ................................................................. 42
Yaw axis control force limits for waveoff (go-around). ............................................................ 42
Yaw axis control force limits for asymmetric thrust during takeoff. ......................................... 42
Yaw axis control force limits with flight control failures. .......................................................... 42
Yaw axis control force limits-control mode change. ................................................................ 42
Yaw axis breakout forces. ....................................................................................................... 42
Flying qualities requirements for the lateral flight path axis .................................................... 42
Dynamic response for lateral translation................................................................................. 42
Flying qualities requirements for combined axes.................................................................... 43
Cross-axis coupling in roll maneuvers. ................................................................................... 43
Crosstalk between pitch and roll controllers. .......................................................................... 43
Control harmony...................................................................................................................... 43
Flight at high angle of attack. .................................................................................................. 43
Warning cues........................................................................................................................... 43
Stalls. ....................................................................................................................................... 43
Stall approach.......................................................................................................................... 43
Stall characteristics.................................................................................................................. 43
Stall prevention and recovery. ................................................................................................. 44
One-engine-out stalls. ............................................................................................................. 44
Post-stall gyrations and spins.................................................................................................. 44
Departure from controlled flight. .............................................................................................. 44
Recovey from post-stall gyrations and spins .......................................................................... 44
Flying qualities requirements in atmospheric disturbances .................................................... 45
Allowable flying qualities degradations in atmospheric disturbances. .................................... 45
Definition of atmospheric disturbance model form.................................................................. 46
Application of disturbance models in analyses. ...................................................................... 46
5.
5.1
5.1.1
VERIFICATION ....................................................................................................................... 47
General requirements-verification ........................................................................................... 47
Loadings-verification. .............................................................................................................. 47
vii
MIL-STD-1797A
CONTENTS
Paragraph
5.1.2
5.1.3
5.1.4
5.1.4.1
5.1.4.2
5.1.4.3
5.1.5
5.1.6
5.1.6.1
5.1.6.2
5.1.6.3
5.1.7
5.1.7.1
5.1.7.2
5.1.7.3
5.1.7.4
5.1.7.5
5.1.7.6
5.1.8
5.1.8.1
5.1.8.2
5.1.9
5.1.10
5.1.11
5.1.11.1
5.1.11.2
5.1.11.3
5.1.11.4
5.1.11.5
5.1.11.6
5.1.11.7
5.1.11.8
5.1.12
5.1.12.1
5.1.12.2
5.1.12.3
5.1.12.4
5.1.12.5
5.1.12.6
5.1.12.7
5.1.12.8
5.1.12.9
5.1.12.10
5.1.12.11
5.1.13
5.1.13.1
5.1.13.2
5.1.13.3
5.1.13.4
5.1.13.5
5.1.13.6
5.2
5.2.1
Page
Moments and products of inertia-verification .......................................................................... 47
Internal and external stores-verification. ................................................................................. 47
Flight Envelopes-verification ................................................................................................... 47
Operational Flight Envelopes-verification................................................................................ 47
Service Flight Envelopes-verification. ..................................................................................... 47
Permissible Flight Envelopes-verification................................................................................ 47
Configurations and States of the aircraft-verification. ............................................................. 47
Aircraft Normal States-verification........................................................................................... 47
Allowable Levels for Aircraft Normal States-verification.......................................................... 47
Flight outside the Service Flight Envelopes-verification.......................................................... 47
Ground operation-verification. ................................................................................................. 47
Aircraft Failure States-verification. .......................................................................................... 47
Allowable Levels for Aircraft Failure States-verification .......................................................... 48
Aircraft Special Failure States-verification............................................................................... 48
Probability calculation-verification. .......................................................................................... 48
Generic failure analysis-verification......................................................................................... 48
When Levels are not specified-verification.............................................................................. 48
Failures outside the Service Flight Envelopes-verification...................................................... 48
Dangerous flight conditions-verification. ................................................................................. 48
Warning and indication-verification. ........................................................................................ 48
Devices for indication, warning, prevention, and recovery-verification. .................................. 48
Interpretation of subjective requirements-verification. ............................................................ 48
Interpretation of quantitative requirements-verification. .......................................................... 48
General flying qualities requirements-verification.................................................................... 48
Buffet-verification..................................................................................................................... 48
Release of stores-verification. ................................................................................................. 48
Effects or armament delivery and special equipment-verification. .......................................... 48
Failures-verification. ................................................................................................................ 48
Control margin-verification....................................................................................................... 48
Pilot-induced oscillations(PIO)-verification. ............................................................................. 49
Residual oscillations-verification. ............................................................................................ 49
Control cross-coupling-verification. ......................................................................................... 49
General flight control system characteristics-verification. ....................................................... 49
Control centering and breakout forces-verification.................................................................. 49
Cockpit control free play-verification. ...................................................................................... 49
Adjustable controls-verification................................................................................................ 49
Rate of control displacement-verification. ............................................................................... 49
Dynamic characteristics-verification. ....................................................................................... 49
Damping-verification................................................................................................................ 49
Transfer to alternate control modes-verification. ..................................................................... 49
Flight control system failures-verification. ............................................................................... 49
Augmentation systems-verification.......................................................................................... 49
Auxiliary dive recovery devices-verification............................................................................. 49
Direct force controllers-verification. ......................................................................................... 49
General trim requirements-verification .................................................................................... 49
Trim system irreversibility-verification. .................................................................................... 49
Rate of trim operation-verification. .......................................................................................... 50
Stalling of trim systems-verification. ........................................................................................ 50
Transients and trim changes-verification................................................................................. 50
Trim for asymmetric thrust-verification. ................................................................................... 50
Automatic trim system-verification........................................................................................... 50
Flying qualities requirements for the pitch axis-verification .................................................... 50
Pitch attitude dynamic response to pitch controller-verification .............................................. 50
viii
MIL-STD-1797A
CONTENTS
Paragraph
5.2.1.1
5.2.1.2
5.2.2
5.2.3
5.2.4
5.2.5
5.2.6
5.2.6.1
5.2.6.2
5.2.7
5.2.7.1
5.2.7.2
5.2.7.2.1
5.2.7.3
5.2.7.4
5.2.8
5.2.8.1
5.2.8.2
5.2.8.3
5.2.8.4
5.2.8.5
5.2.8.6
5.2.8.6.1
5.2.8.7
5.2.8.7.1
5.2.8.7.2
5.2.8.7.3
5.2.8.7.4
5.2.8.8
5.2.9
5.2.9.1
5.2.9.2
5.3
5.3.1
5.3.1.1
5.3.1.2
5.3.2
5.3.3
5.3.3.1
5.3.3.2
5.3.4
5.4
5.4.1
5.4.1.1
5.5
5.5.1
5.5.1.1
5.5.1.2
5.5.1.3
5.5.1.4
5.5.1.5
5.5.2
5.5.3
Page
Long-term pitch response-verification. .................................................................................... 50
Short-term pitch response-verification..................................................................................... 50
Pilot-induced pitch oscillations-verification.............................................................................. 50
Residual pitch oscillations-verification. ................................................................................... 50
Normal acceleration at pilot station-verification. ..................................................................... 50
Pitch trim changes-verification. ............................................................................................... 50
Pitch axis response to other inputs-verification....................................................................... 50
Pitch axis response to failures, controls free-verification. ....................................................... 50
Pitch axis response to configuration or control mode change-verification. ............................. 50
Pitch axis control power-verification........................................................................................ 50
Pitch axis control power in unaccelerated flight-verification.................................................... 50
Pitch axis control power in maneuvering flight-verification...................................................... 50
Load factor response-verification. ........................................................................................... 51
Pitch axis control power in takeoff-verification. ....................................................................... 51
Pitch axis control power in landing-verification........................................................................ 51
Pitch axis control forces-verification........................................................................................ 51
Pitch axis control forces-steady-state control force per g-verification. .................................... 51
Pitch axis control forces-transient control force per g-verification........................................... 51
Pitch axis control forces-control force variations during rapid speed changes-verification..... 51
Pitch axis control forces-control force vs. control deflection-verification................................. 51
Pitch axis control breakout forces-verification. ........................................................................ 51
Pitch axis control force limits-verification................................................................................. 51
Pitch axis control force limits-takeoff-verification..................................................................... 51
Pitch axis control force limits-landing-verification.................................................................... 51
Pitch axis control force limits-dives-verification. ...................................................................... 51
Pitch axis control force limits-sideslips-verification.................................................................. 51
Pitch axis control force limits-failures-verification.................................................................... 51
Pitch axis control force limits-control mode change-verification.............................................. 51
Pitch axis trim systems-verification. ........................................................................................ 51
Pitch axis control displacements-verification .......................................................................... 52
Pitch axis control displacements-takeoff-verification............................................................... 52
Pitch axis control displacements-maneuvering-verification..................................................... 52
Flying qualities requirements for the normal (flight path) axis-verification .............................. 52
Flight path response to attitude change-verification ............................................................... 52
Transient flight path response to attitude change-verification................................................. 52
Steady-state flight path response to attitude change-verification............................................ 52
Flight path response to designated flight path controller-verification...................................... 52
Flight path control power-verification ...................................................................................... 52
Control power for designated primary flight path controller-verification. ................................. 52
Control power for designated secondary flight path controller-verification. ............................ 52
Flight path controller characteristics-verification. .................................................................... 52
Flying qualities requirements for the longitudinal (speed) axis-verification ............................ 52
Speed response to attitude changes-verification. ................................................................... 52
Speed response to attitude changes-relaxation in transonic flight-verification. ...................... 52
Flying qualities requirements for the roll axis-verification ....................................................... 52
Roll response to roll controller-verification .............................................................................. 52
Roll mode-verification. ............................................................................................................. 52
Spiral stability-verification........................................................................................................ 52
Coupled roll-spiral oscillation-verification................................................................................ 52
Roll oscillations-verification. .................................................................................................... 52
Roll time delay-verification....................................................................................................... 52
Pilot-induced roll oscillations-verification. ............................................................................... 53
Linearity of roll response to roll controller-verification............................................................. 53
ix
MIL-STD-1797A
CONTENTS
Paragraph
5.5.4
5.5.5
5.5.6
5.5.7.1
5.5.7.2
5.5.8
5.5.8.1
5.5.8.2
5.5.8.3
5.5.8.4
5.5.8.5
5.5.8.6
5.5.9
5.5.9.1
5.5.9.2
5.5.9.3
5.5.9.4
5.5.9.5
5.5.9.5.1
5.5.9.5.2
5.5.9.5.3
5.5.9.5.4
5.5.9.5.5
5.5.9.5.6
5.5.9.5.7
5.6
5.6.1
5.6.1.1
5.6.1.2
5.6.1.3
5.6.2
5.6.3
5.6.4
5.6.5
5.6.5.1
5.6.5.2
5.6.5.3
5.6.6
5.6.6.1
5.6.6.2
5.6.6.3
5.6.7
5.6.7.1
5.6.7.2
5.6.7.3
5.6.7.4
5.6.7.5
5.6.7.6
5.6.7.7
5.6.7.8
5.6.7.9
5.6.7.10
5.6.7.11
Page
Lateral acceleration at the pilot station-verification. ................................................................ 53
Roll characteristics in steady sideslip-verification. .................................................................. 53
Roll axis control for takeoff and landing in crosswinds-verification......................................... 53
Roll axis response to augmentation failures-verification. ........................................................ 53
Roll axis response to configuration or control mode change-verification. .............................. 53
Roll axis control power-verification ......................................................................................... 53
Roll axis response to augmentation failures-verification. ........................................................ 53
Roll axis control power in steady slips-verification. ................................................................. 53
Roll axis control power in crosswinds-verification................................................................... 53
Roll axis control power for asymmetric thrust-verification....................................................... 53
Roll axis control power in dives and pullouts-verification........................................................ 53
Roll axis control power for asymmetric loading-verification. ................................................... 53
Roll axis control forces and displacements-verification........................................................... 54
Roll control displacements-verification. ................................................................................... 54
Roll axis control forces to achieve required roll performance-verification............................... 54
Roll axis control sensitivity-verification. ................................................................................... 54
Roll axis control centering and breakout forces-verification.................................................... 54
Roll axis control force limits-verification .................................................................................. 54
Roll axis control force limits in steady turns-verification.......................................................... 54
Roll axis control force limits in dives and pullouts-verification................................................. 54
Roll axis control force limits in crosswinds-verification............................................................ 54
Roll axis control force limits in steady sideslips-verification. ................................................... 54
Roll axis control force limits for asymmetric thrust-verification................................................ 54
Roll axis control force limits for failures-verification................................................................. 54
Roll axis control force limits for configuration or control mode change-verification. ............... 54
Flying qualities requirements for the yaw axis-verification...................................................... 54
Yaw axis response to yaw and side-force controllers-verification .......................................... 54
Dynamic lateral-directional response-verification.................................................................... 54
Steady sideslips-verification. ................................................................................................... 54
Wings-level turn-verification. ................................................................................................... 54
Yaw axis response to roll controller-verification...................................................................... 54
Pilot-induced yaw oscillations-verification............................................................................... 55
Yaw axis control for takeoff and landing in crosswinds-verification. ....................................... 55
Yaw axis response to other inputs-verification........................................................................ 55
Yaw axis response to asymmetric thrust-verification. ............................................................. 55
Yaw axis response to failures-verification. .............................................................................. 55
Yaw axis response to configuration or control mode change-verification. .............................. 55
Yaw axis control power-verification......................................................................................... 55
Yaw axis control power for takeoff, landing, and taxi-verification............................................ 55
Yaw axis control power for asymmetric thrust-verification. ..................................................... 55
Yaw axis control power with asymmetric loading-verification. ................................................ 55
Yaw axis control forces-verification......................................................................................... 55
Yaw axis control force limits in rolling maneuvers-verification. ............................................... 55
Yaw axis control force limits in steady turns-verification. ........................................................ 55
Yaw axis control force limits during speed changes-verification. ............................................ 55
Yaw axis control force limits in crosswinds-verification........................................................... 55
Yaw axis control force limits with asymmetric loading-verification. ......................................... 55
Yaw axis control force limits in dives and pullouts-verification................................................ 55
Yaw axis control force limits for waveoff (go-around)-verification. .......................................... 55
Yaw axis control force limits for asymmetric thrust during takeoff-verification........................ 56
Yaw axis control force limits with flight control failures-verification. ........................................ 56
Yaw axis control force limits-control mode change-verification............................................... 56
Yaw axis breakout forces-verification...................................................................................... 56
x
MIL-STD-1797A
CONTENTS
Paragraph
Page
5.7
5.7.1
5.8
5.8.1
5.8.2
5.8.3
5.8.4
5.8.4.1
5.8.4.2
5.8.4.2.1
5.8.4.2.2
5.8.4.2.3
5.8.4.2.4
5.8.4.3
5.8.4.3.1
5.8.4.3.2
5.9
5.9.1
5.9.2
5.9.3
Flying qualities requirements for the lateral flight path axis-verification.................................. 56
Dynamic response for lateral translation-verification. ............................................................. 56
Flying qualities requirements for combined axes-verification ................................................. 56
Cross-axis coupling in roll maneuvers-verification.................................................................. 56
Crosstalk between pitch and roll controllers-verification......................................................... 56
Control harmony-verification. .................................................................................................. 56
Flight at high angle of attack-verification ................................................................................. 56
Warning cues-verification. ....................................................................................................... 56
Stalls-verification. .................................................................................................................... 56
Stall approach-verification. ...................................................................................................... 56
Stall characteristics-verification. .............................................................................................. 56
Stall prevention and recovery-verification. .............................................................................. 56
One-engine-out stalls-verification............................................................................................ 56
Post-stall gyrations and spins-verification. .............................................................................. 56
Departure from controlled flight-verification............................................................................. 56
Recovery from post-stall gyrations and spins-verification. ...................................................... 57
Flying qualities requirements in atmospheric disturbances-verification.................................. 57
Allowable flying qualities degradations in atmospheric disturbances-verification................... 57
Definition of atmospheric disturbance model form-verification. .............................................. 57
Application of disturbance models in analyses-verification..................................................... 57
6.
6.1
6.2
6.3
6.4
6.5
6.6
6.7
NOTES .................................................................................................................................... 57
Intended use............................................................................................................................ 57
Level definitions. ..................................................................................................................... 57
Reference documents tree...................................................................................................... 57
Data requirements................................................................................................................... 58
Subject term (key word) listing ................................................................................................ 59
Responsible engineering office (REO).................................................................................... 59
Changes from previous issue.................................................................................................. 59
FIGURES
Figure
Page
Figure 1. Roll-sideslip coupling parameters--right rolls......................................................................... 10/11
Figure 2. Roll-sideslip coupling parameters--left rolls........................................................................... 12/13
TABLES
Table
TABLE I.
TABLE II.
TABLE III.
TABLE IV.
TABLE V.
TABLE VI.
Page
Operational Flight Envelope. ................................................................................................... 20
Aircraft Normal States. ............................................................................................................ 22
Levels for Aircraft Failure States. ............................................................................................ 23
Pitch trim change conditions.................................................................................................... 30
Flying qualities in atmospheric disturbances for Aircraft Normal States. ................................ 45
Flying qualities in atmospheric disturbances for Aircraft Failure States.................................. 46
xi
MIL-STD-1797A
1.
SCOPE
1.1
Purpose.
This standard contains the requirements for the flying and ground handling qualities of
_______________. It is intended to assure flying qualities for adequate mission performance and flight
safety regardless of the design implementation or flight control system augmentation.
1.2
Applicability.
The requirements of this standard, with blanks filled in, are to be applied during the design, construction,
testing and acceptance of the subject aircraft.
2.
APPLICABLE DOCUMENTS
2.1
Government documents.
2.1.1 Specifications, standards, and handbooks.
The following specifications, standards, and handbooks form a part of this document to the extent
specified herein. Unless otherwise specified, the issues of these documents are those listed in the issue
of the Department of Defense Index of Specifications and Standards (DODISS) and supplement thereto,
cited in the solicitation (see 6.2).
SPECIFICATIONS
Military
MIL-C-18244
Control and Stabilization Systems: Automatic, Piloted Aircraft, General
Specification for
MIL-F-87242
Flight Control Systems, General Specification for
(Unless otherwise indicated, copies of federal and military specifications, standards and handbooks are
available from Military Specifications and Standards, Bldg 4D, 700 Robbins Avenue, Philadelphia, PA
19111-5094).
2.1.2 Other Government documents, drawings, and publications.
The following other Government documents, drawings, and publications form a part of this standard to the
extent specified herein. Unless otherwise specified, the issues are those cited in the solicitation.
2.2
Non-Government publications.
The following document(s) form a part of this standard to the extent specified herein. Unless otherwise
specified, the issues of the documents which are DoD adopted are those listed in the issue of the
DODISS cited in the solicitation. Unless otherwise specified, the issues of the documents not listed in the
DODISS are the issues of the documents cited in the solicitation.
(Application for copies should be addressed to ______________.)
(Nongovemment standards and other publications are normally available from organizations that prepare
or distribute the documents. These documents also may be available in or through libraries or other
information services.)
2.3
Order of precedence.
In the event of a conflict between the text of this document and the references cited herein, the text of this
document shall take precedence. Nothing in this document, however, supersedes applicable laws and
regulations unless a specific exemption has been obtained.
1
MIL-STD-1797A
3.
DEFINITIONS
3.1
Aircraft classification and operational missions.
For the purpose of this standard, the aircraft specified in this requirement is to accomplish the following
missions: __________. The aircraft thus specified will be a Class ______ aircraft. The letter -L following a
class designation identifies an aircraft as land-based; carrier-based aircraft are similarly identified by -C.
When no such differentiation is made in a requirement, the requirement applies to both land-based and
carrier-based aircraft.
3.2
Flight Phase Categories.
To accomplish the mission requirements the following general Flight Phase Categories are involved:
_______. Special Flight Phases to be considered are: ________.
3.3
Levels and qualitative suitability of flying qualities.
The handling characteristics described in this standard are specified in terms of qualitative degrees of
suitability and Levels. The degrees of suitability are defined as:
Satisfactory
Flying qualities clearly adequate for the mission Flight Phase. Desired
performance is achievable with no more than minimal pilot compensation.
Acceptable
Flying qualities adequate to accomplish the mission Flight Phase, but some
increase in pilot workload or degradation in mission effectiveness, or both, exists.
Controllable
Flying qualities such that the aircraft can be controlled in the context of the
mission Flight Phase, even though pilot workload is excessive or mission
effectiveness is inadequate, or both. The pilot can transition from Category A
Flight Phase tasks to Category B or C Flight Phases, and Category B and C
Flight Phase tasks can be completed.
Level 1 is Satisfactory, Level 2 is Acceptable, and Level 3 is Controllable. In the presence of higher
intensities of atmospheric disturbances, 4.9.1 states the relationship between Levels and qualitative
degrees of suitability. Where possible, the flying qualities requirements are stated for each Level in terms
of limiting values of one or more parameters. Each value, or combination of values, represents a
minimum condition necessary to meet one of the three Levels of acceptability.
It is to be noted that Level 3 is not necessarily defined as safe. This is consistent with the Cooper-Harper
rating scale: for Cooper-Harper ratings of 8 and 9, controllability may be in question. If safe characteristics
are required for Level 3, then action must be taken to improve aircraft flying qualities.
In some cases sufficient data do not exist to allow the specification of numerical values of a flying quality
parameter. In such cases it is not possible to define explicitly a quantitative boundary of each Level, so
the required Levels are then to be interpreted in terms of qualitative degrees of suitability for the piloting
tasks appropriate for mission accomplishment.
3.4
Parameters.
Terms and symbols used throughout this standard are defined as follows:
3.4.1
S
General terms
Wing area
s
Laplace operator
q
Dynamic pressure
MSL
Mean Sea Level
T2
Time
to
double
amplitude;
T2
=
-0.693/ζωn
for
oscillations,
T2 = -0.693T for first-order divergencies, T is the modal time constant
T
Time delay
2
MIL-STD-1797A
Service ceiling
Altitude at a given airspeed at which the rate of climb is 100 ft/min at the stated
weight and engine thrust
Combat ceiling
Altitude at a given airspeed at which the rate of climb is 500 ft/min at the stated
weight and engine thrust
Cruising ceiling
Altitude at a given airspeed at which the rate of climb is 300 ft/min at NRT at the
stated weight
hmax
Maximum service altitude (defined in 4.1.4.2)
ho
Maximum operational altitude (4.1.4.1)
homin
Minimum operational altitude (4.1.4.1)
c.g.
Aircraft center of gravity
3.4.2 Speeds
Airspeed
Magnitude of the velocity with respect to the air mass
max
Equivalent airspeed,
EAS
True airspeed multiplied by σ where σ, is the ratio of free-stream density at
the given altitude to standard sea-level air density
Calibrated airspeed,
CAS
Airspeed-indicator reading corrected for position and instrument error but not for
compressibility
Refusal speed
The maximum speed to which the aircraft can accelerate and then stop in the
available runway length
M
Mach number
V
Airspeed along the flight path (where appropriate, V may be replaced by M in this
standard)
VS
Stall speed (equivalent airspeed), at 1g normal to the flight path, defined as the
highest of:
a. Speed for steady straight flight at CLmax, the first local maximum of the curve of
lift coefficient (L/ q S) vs. angle of attack which occurs as CL is increased from
zero.
b. Speed at which uncommanded pitching, rolling or yawing occurs (4.8.4.2).
c. Speed at which intolerable buffet or structural vibration is encountered.
3
MIL-STD-1797A
Conditions for determining VS
The aircraft shall be initially trimmed at approximately 1.2 VS with the following settings, after which the
trim and throttle settings shall be held constant:
FLIGHT PHASE
THRUST SETTINGS*
TRIM SETTING
Climb (CL)
Normal climb
For straight flight
Descent (D)
Normal descent
For straight flight
Emergency descent (ED)
Idle
For straight flight
Emergency deceleration (DE)
Idle
For straight flight
Takeoff (TO)
Takeoff
Recommended takeoff setting
Approach (PA)
Normal approach
For normal approach
Waveoff/Go-around (WO)
Takeoff
For normal approach
Landing (L)
Idle
For normal approach
All other
TLF at 1.2 VS
For straight flight
* - Either on all engines or on remaining engines with critical engine inoperative, whichever yields the
higher value of VS.
In flight test, it is necessary to reduce speed very slowly (typically 1/2 knot per second or less) to minimize
dynamic lift effects. The load factor will generally not be exactly 1g when stall occurs; when this is the
case, Vs is defined as follows:
V
VS =
nf
where V and nf are the measured values at stall, nf being the load factor normal to the flight path.
VS(X), Vmin (X), Vmax(X)
Short-hand notation for the speeds VS, Vmin, Vmax for a given configuration,
weight, center-of-gravity position, and external store combination associated
with Flight Phase X. For example, the designation Vmax(TO) is used in 4.2.8.6.1
to emphasize that the speed intended (for the weight, center of gravity and
external store combination under consideration) is Vmax for the configuration
associated with the takeoff Flight Phase. This is necessary to avoid confusion,
since the configuration and Flight Phase change from takeoff to climb during
the maneuver.
Vtrim
Trim speed
4
MIL-STD-1797A
Vend
Speed for maximum endurance
VL/D
Speed for maximum lift-to-drag ratio
VR/C
Speed for maximum rate of climb
Vrange
Speed for maximum range with zero wind
VNRT
High speed, level flight, normal rated thrust
VMRT
High speed, level flight, military rated thrust
VMAT
High speed, level flight, maximum augmented thrust
Vmax
Maximum service speed (defined in 4.1.4.2)
Vmin
Minimum service speed (defined in 4.1.4.2)
Vo
Maximum operational speed (4.1.4.1)
Vo
Minimum operational speed (4.1.4.1)
VG
Gust penetration speed
VMCA, VMC
Minimum controllable airspeed while airborne
VMCG
Minimum controllable airspeed while on the ground
max
min
3.4.3 Thrust and power
Thrust and power
For propeller-driven aircraft, the word “thrust" shall be replaced by the word
“power" throughout the standard
TLF
Thrust for level flight
NRT
Normal rated thrust, which is the maximum thrust at which the engine can be
operated continuously
MRT
Military rated thrust, which is the maximum thrust at which the engine can be
operated for a specified period
MAT
Maximum augmented thrust:
available for the Flight Phase
Takeoff thrust
Maximum thrust available for takeoff
5
maximum thrust, augmented by all means
MIL-STD-1797A
3.4.4 Control parameters
Pitch, roll, yaw
The stick or wheel and pedals manipulated by the pilot to produce pitching,
controls
rolling and yawing moments respectively; the cockpit controls
Pitch control force,
Fes
Component of applied force, exerted by the pilot on the cockpit control in the
direction to command pitch, acting at the center of the stick grip or wheel in a
direction perpendicular to a line between the center of the stick grip and the
stick control column pivot
δes
Deflection of the cockpit pitch controller
Roll control force,
For a stick control, the component of control force exerted by the pilot in the
direction to command roll, acting at the center of the stick grip in a direction
perpendicular to a line between the center of the stick grip and the stick pivot.
For a wheel control, the total moment applied by the pilot about the wheel axis
in the plane of the wheel divided by the average radius from the wheel pivot to
the pilot's grip
Fas
δas
Deflection of the cockpit roll controller
Yaw-control pedal
force,
Frp
Difference of push-force components of forces exerted by the pilot on the yawcontrol pedals, lying in planes parallel to the plane of symmetry, measured
perpendicular to the pedals at the normal point of application of the pilot's
instep on the respective yaw-control pedals.
δrp
Deflection of the cockpit yaw controller
Control effectors
The aerodynamic surfaces, thrust vectoring, reaction controls, etc., which
produce forces and moments to control the aircraft.
Direct normal force
control
A device producing direct normal force for the primary purpose of controlling
the flight path of the aircraft. The term describes the concept of directly
modulating the normal force on an aircraft by changing its lifting capabilities at
a constant angle of attack and constant airspeed or by controlling the normal
force component of such items as jet exhausts, propellers, and fans
Control power
Effectiveness of control surfaces in applying forces or moments to an aircraft.
For example, 50 percent of available roll control power is 50 percent of the
maximum rolling moment that is available to the pilot with allowable roll control
force
3.4.5
1/Tθl
Longitudinal parameters
Low-frequency pitch attitude zero
1/Tθ2
High-frequency pitch attitude zero
ζsp
Damping ratio of the short-period oscillation
6
MIL-STD-1797A
ωsp
Undamped natural frequency of the short-period oscillation
ζp
Damping ratio of the phugoid oscillation
ωp
Undamped natural frequency of the phugoid oscillation
ωBW
Bandwidth frequency, with stated minimum phase and gain margins (generally
60 deg and 6 dB resp.)
n
Normal acceleration or normal load factor, g's, measured at the c.g. unless the
location is otherwise specified
nz
Normal acceleration measured at the instantaneous center of rotation for pitch
control inputs
nL
Symmetrical flight limit load factor for a given Aircraft Normal State, based on
structural considerations
nmax, nmin
Maximum and minimum service load factors
n (+), n (-)
For a given altitude, the upper and lower boundaries of n in the V-n diagrams
depicting the Service Flight Envelope
nomax, nomin
Maximum and minimum operational load factors
no (+), no (-)
For given altitude, the upper and lower boundaries of n in the V-n diagrams
depicting the Operational Flight Envelope (4.1.4.1)
α
Angle of attack; the angle in the plane of symmetry between the fuselage
reference line and the tangent to the flight path at the aircraft center of gravity
αS
The stall angle of attack at constant speed for the configuration, weight, centerof-gravity position and external store combination associated with a given Aircraft
Normal State; defined as the lowest of the following:
a. Angle of attack for the highest steady load factor, normal to the flight path,
that can be attained at a given speed or Mach number
b. Angle of attack, for a given speed or Mach number, at which uncommanded
pitching, rolling or yawing occurs (4.8.4.2)
c. Angle of attack, for a given speed or Mach number, at which intolerable
buffeting is encountered
CLstall
Lift coefficient at αS defined above
n/α
The steady-state normal acceleration change per unit change in angle of attack
for an incremental pitch control deflection at constant speed (airspeed and Mach
number)
7
MIL-STD-1797A
Fs/n
Gradient of steady-state pitch control force versus n at constant speed (4.2.8.1)
γ
Climb angle, positive for climbing flight
y = sin-1 (vertical speed/true airspeed)
θ
Pitch attitude, the angle between the x-axis and the horizontal
L
Aerodynamic lift plus thrust component normal to the flight path
3.4.6
TR
Lateral-directional parameters
First-order roll mode time constant, positive for stable mode
TS
First-order spiral mode time constant, positive for stable mode
ωφ
Undamped natural frequency of numerator quadratic of φ/Fas transfer function
ζφ
Damping ratio of numerator quadratic of φ/Fas transfer function
ωd
Undamped natural frequency of the dutch roll oscillation
ζd
Damping ratio of the dutch roll oscillation
Td
Damped period of the dutch roll,
Td =
2π
ω d 1 − ζ 2d
ωRS
Undamped natural frequency of a coupled roll-spiral oscillation
ζRS
Damping ratio of a coupled roll-spiral oscillation
φ
Bank angle measured in the y-z plane, between the y-axis and the horizontal
φt
Bank angle change in time t, in response to a control deflection of the form given
in 4.5.8.1
p
Roll rate about the x-axis
δ'rp(3)
Yaw-to-roll crossfeed parameter (4.6.2)
posc/pav
A measure of the ratio of the oscillatory component of roll rate to the average
component of roll rate following a yaw-control-free step roll control command:
p
ζ d ≤ 0.2 : osc =
p av
p1 + p 3 − 2p 2
p1 + p 3 + 2p 2
p
ζ d > 0.2 : osc =
p av
p1 − p 2
p1 + p 2
8
MIL-STD-1797A
where p1, p2 and p3 are roll rates at the first, second and third peaks, respectively
(figures 1 and 2)
φosc/φav
A measure of the ratio of the oscillatory component of bank angle to the average
component of bank angle following a pedals-free impulse aileron control
command:
ζ d ≤ 0 .2 :
φ osc
φ + φ 3 − 2φ 2
= 1
φ av
φ1 + φ 3 + 2φ 2
φ
φ − φ2
ζ d > 0.2 : osc = 1
φ av
φ1 + φ 2
where φ1, φ2, and φ3 are bank angles at the first, second and third peaks,
respectively
β
Sideslip angle at the center of gravity, angle between undisturbed flow and plane
of symmetry; positive, or right sideslip corresponds to incident flow approaching
from the right side of the plane of symmetry
∆βmax
Maximum change in sideslip occurring within 2 seconds or one half-period of the
dutch roll, whichever is greater, for a step roll-control command (figures 1 and 2)
9
MIL-STD-1797A
Figure 1. Roll-sideslip coupling parameters--right rolls.
10
MIL-STD-1797A
FIGURE 1. Roll-sideslip coupling parameters--right rolls. -continued
11
MIL-STD-1797A
Figure 2. Roll-sideslip coupling parameters--left rolls.
12
MIL-STD-1797A
FIGURE 2. Roll-Sideslip coupling parameters--left rolls. - continued
13
MIL-STD-1797A
k
Ratio of "commanded roll performance” to “applicable roll performance
requirement" of 4.5.8.1, where:
a. “Applicable roll performance requirement," (φt)requirement is determined from
4.5.8.1 for the Class and Flight Phase Category under consideration
b. “Commanded roll performance," ( φt)command, is the bank angle attained in the
stated time for a given step roll command with yaw control pedals employed as
specified in 4.5.8.1
k=
(φt )common
(φt )requirement
tnβ
Time for the dutch roll oscillation in the sideslip response to reach the nth local
maximum for a right step or pulse roll-control command, or the nth local minimum
for a left command. In the event a step control input cannot be accomplished, the
control shall be moved as abruptly as practical and, for purposes of this
definition, time shall be measured from the instant the cockpit control deflection
passes through half the amplitude of the commanded value. For pulse inputs,
time shall be measured from a point halfway through the duration of the pulse
ψβ
Phase angle expressed as a lag for a cosine representation of the dutch roll
oscillation in sideslip, where
ψβ = −
360
t n + (n − 1) 360 (deg rees )
Td β
with n as used in the definition of tnβ above
∠p / β
Phase angle between roll rate and sideslip in the free dutch roll oscillation. Angle
is positive when p leads 0 by an angle between 0 and 180 deg
φ/β d
At any instant, the ratio of amplitudes of the bank-angle and sideslip-angle
envelopes in the dutch roll mode
14
MIL-STD-1797A
Examples showing measurement of roll-sideslip coupling parameters are shown on figure 1 for right rolls
and figure 2 for left rolls. Since several oscillations of the dutch roll are required to measure these
parameters, and since for proper identification large roll rates and bank angle changes must generally be
avoided, step roll control inputs should be small. It should be noted that since ψβ is the phase angle of the
dutch roll component of sideslip, care must be taken to select a peak far enough downstream that the
position of the peak is not influenced by the roll mode. In practice, peaks occurring one or two roll mode
time constants after the aileron input will be relatively undistorted. Care must also be taken when there is
ramping of the sideslip trace, since ramping will displace the position of a peak of the trace from the
corresponding peak of the dutch roll component. In practice, the peaks of the dutch roll component of
sideslip are located by first drawing a line through the ramping portion of the sideslip trace and then
noting the times at which the vertical distance between the line and the sideslip trace is the greatest [See
Case (a) of the following enlarged section of figures 1 and 2].
Since the first local maximum of the dutch roll component of the sideslip response occurs at t = 2.95
seconds,
ψβ = −
−360
360
(2.95) = −303 deg rees
t n + (n − 1) 360 (deg rees ) =
3.5
Td β
15
MIL-STD-1797A
For example, if the roll performance requirement is φt = 30 degrees in one second with rudder pedals free
(as in the rolls of 4.5.8.1), then from the definitions, “k" for this condition is
k=
(φt )common
(φt )requirement
Therefore from figures 1 and 2:
Case (a): k = 9.1/30 = 0.30
Case (c): k
6.8/30
Case (b): k = 8.1/30 = 0.27
Case (d): k
6.0/30
0.23
0.20
r&o / p& o
Ratio of initial yawing acceleration to initial rolling acceleration for a step roll
control input (equivalent to N′δas L′δas )
µ
Rudder shaping parameter:
µ = δrp (3) -1 where δrp (3) is the rudder pedal deflection, 3 seconds after a step
roll control input, that would be required for perfect coordination (see 4.6.2
Guidance)
3.4.7
j
Atmospheric disturbance parameters
−1
Ω
Spatial (reduced) frequency (radians per foot)
ω
Temporal frequency (radians per second), where ω = ΩV
t
Time (seconds)
ug
Translational disturbance velocity along the x-axis, positive forward (feet per
second)
vg
Translational disturbance velocity along the y-axis, positive to pilot's right (feet
per second)
wg
Translational disturbance velocity along the z-axis, positive down (feet per
second)
Note: Random ug, vg, wg have Gaussian (normal) distributions
Vw/d
Velocity of wind over deck (feet per second)
16
MIL-STD-1797A
σ, RMS
Root-mean-square disturbance intensity, where
∞
∞
0
0
σ 2 = ∫ Φ (Ω ) dΩ = ∫ φ(ω) dω
σu
Root-mean-square intensity of ug (feet per second)
σv
Root-mean-square intensity of vg (feet per second)
σw
Root-mean-square intensity of wg (feet per second)
Lu
Scale for ug (feet)
Lv
Scale for vg (feet)
Lw
Scale for wg (feet)
Φug(Ω)
Spectrum for ug, where Φug(Ω) = Vφug(ω) (ft3/sec2)
Φvg(Ω)
Spectrum for vg, where Φvg(Ω) = Vφvg(ω) (ft3/sec2)
Φwg(Ω)
Spectrum for wg, where Φwg(Ω) = Vφwg(ω) (ft3/sec2)
vm
Generalized discrete gust intensity, positive along the positive axes, m = x, y, z
(feet per second)
dm
Generalized discrete gust length (always positive), m = x, y, z (feet)
u20
Wind speed at 20 feet above the ground (feet per second)
X
Distance from aircraft to ship center of pitch, negative aft of ship (feet)
ψw
Mean wind direction relative to runway
pg
Rotary disturbance velocity about the x-axis (radians per second)
qg
Rotary disturbance velocity about the y-axis (radians per second)
rg
Rotary disturbance velocity about the z-axis (radians per second)
Φpg
Spectrum for pg (ft/sec2)
Φqg
Spectrum for qg (ft/sec2)
Φ rg
Spectrum for rg (ft/sec2)
17
MIL-STD-1797A
3.5
Terms used in high angle of attack requirements
Post-stall
The flight regime involving angles of attack greater than nominal stall angles of
attack. The aircraft characteristics in the post-stall regime may consist of three
more or less distinct consecutive types of aircraft motion following departure
from controlled flight: post-stall gyration, incipient spin, and developed spin
Post-stall gyration
(PSG)
Uncontrolled motions about one or more aircraft axes following departure from
controlled flight. While this type of aircraft motion involves angles of attack
higher than stall angle, lower angles may be encountered intermittently in the
course of the motion
Spin
That part of the post-stall aircraft motion which is characterized by a sustained
yaw rotation. The spin may be upright or inverted, flat (high angle of attack) or
steep (low but still stalled angle of attack) and the rotary motions may have
oscillations in pitch, roll and yaw superimposed on them. The incipient spin is
the initial, transient phase of the motion during which it is not possible to
identify the spin mode, usually followed by the developed spin, the phase
during which it is possible to identify the spin mode.
18
MIL-STD-1797A
4.
REQUIREMENTS
4.1
General requirements
4.1.1 Loadings.
The contractor shall define the longitudinal, lateral, and vertical envelopes of center of gravity and corresponding
weights that will exist for each Flight Phase. Throughout these envelopes shall include the most forward and aft
center-of-gravity positions as defined in ___________. In addition the contractor shall determine the maximum
center-of-gravity excursions attainable through failures in systems or components, such as fuel sequencing or
hung stores, for each Flight Phase. Throughout these envelopes, plus a growth margin of ________, and for the
excursions cited, this standard applies.
4.1.2 Moments and products of inertia.
The contractor shall define the moments and products of inertia of the aircraft associated with all loadings of
4.1.1. The requirements of this standard shall apply for all moments and products of inertia so defined.
4.1.3 Internal and external stores.
The symmetric and asymmetric store combinations to be considered are as follows: _________. The
requirements of this standard shall apply to these store conditions. The effects of stores on the weight, moments
of inertia, center-of-gravity position, and aerodynamic characteristics of the aircraft shall be determined for each
mission Flight Phase. When the stores contain expendable loads, the requirements of this standard apply
throughout the range of store loadings, including sloshing/shifting.
4.1.4
Flight Envelopes
4.1.4.1
Operational Flight Envelopes.
The Operational Flight Envelopes define the boundaries in terms of speed, altitude and load factor within which
the aircraft must be capable of operating in order to accomplish the missions of 3.1 and in which Level 1 flying
qualities are required. These envelopes shall implicitly include the ranges of other parameters, such as sideslip,
which may normally be encountered. The range of sideslip or lateral acceleration employed with direct side force
control is to be stated explicitly. In the absence of other specific instructions, the contractor shall use the
representative conditions of table I for the applicable Flight Phases.
4.1.4.2
Service Flight Envelopes.
For each Aircraft Normal State the contractor shall establish, subject to the approval of the procuring activity,
Service Flight Envelopes showing combinations of speed, altitude, and normal acceleration derived from aircraft
limits as distinguished from mission requirements. These envelopes shall implicitly include the ranges of other
parameters, such as sideslip, which can be expected within the speed, altitude and load-factor bounds. For each
applicable Flight Phase and Aircraft Normal State, the boundaries of the Service Flight Envelopes can be
coincident with or lie outside the corresponding Operational boundaries.
4.1.4.3
Permissible Flight Envelopes.
The contractor shall define Permissible Flight Envelopes, subject to the approval of the procuring activity, which
encompass all regions in which operation of the aircraft is both allowable and possible, and which the aircraft is
capable of safely encountering. These Envelopes define boundaries in terms of speed, altitude, load factor and
any other flight limits.
4.1.5 Configurations and States of the aircraft.
The requirements of this standard apply for all configurations required or encountered in the applicable Flight
Phases of 3.2. A selected configuration is defined by the positions and adjustments of the various selectors and
controls available to the crew except for pitch, roll, yaw, throttle and trim controls. Examples are: the flap control
setting and the yaw damper ON or OFF. The selected configurations to be examined must include those required
for performance demonstration and mission accomplishment. Additional configurations to be investigated are
defined as follows: ______. Switches which activate stability augmentation necessary to meet the requirements of
this standard are considered always to be ON unless otherwise specified. The State of the aircraft is defined by
19
MIL-STD-1797A
TABLE I. Operational Flight Envelope.
FLIGHT
PHASE
CATEGORY
AIRSPEED
FLIGHT PHASE
A
B
C
20
ALTITUDE
LOAD FACTOR
MIL-STD-1797A
the selected configuration together with the functional status of each of the aircraft components or systems,
throttle setting, weight, moments of inertia, center-of-gravity position, and external store complement. The trim
setting and the positions of the pitch, roll, and yaw controls are not included in the definition of Aircraft State since
they are often specified in the requirements.
4.1.6 Aircraft Normal States.
The contractor shall define and tabulate all pertinent items to describe the Aircraft Normal States (no component
or system failure) associated with each of the applicable Flight Phases. This tabulation shall be in the format and
use the nomenclature of table II. Certain items, such as weight, moments of inertia, center-of-gravity position,
wing sweep, or thrust setting may vary continuously over a range of values during a Flight Phase. The contractor
shall replace this continuous variation by a limited number of values of the parameter in question which will be
treated as specific States, and which include the most critical values and the extremes encountered during the
Flight Phase in question.
4.1.6.1
Allowable Levels for Aircraft Normal States.
Flying qualities for Aircraft Normal States within the Operational Flight Envelope shall be Level 1 - Flying qualities
for Aircraft Normal States within the Service Flight Envelope but outside the Operational Flight Envelope shall be
Level 2 or better. To account for the natural degradation of pilot-vehicle performance and workload in intense
atmospheric disturbances, the requirements of 4.1.6.1 through 4.1.6.3 are adjusted according to 4.9.1.
4.1.6.2
Flight outside the Service Flight Envelopes.
From all points in the Permissible Flight Envelopes and outside the Service Flight Envelopes, it shall be possible
readily and safely to return to the Service Flight Envelope without exceptional pilot skill or technique. The
requirements on flight at high angle of attack, dive characteristics, dive recovery devices and dangerous flight
conditions shall also apply in all pertinent parts of the Permissible Flight Envelopes.
4.1.6.3
Ground operation.
Some requirements pertaining to taxiing, takeoffs and landing involve operation outside the Operational, Service,
and Permissible Flight Envelopes, as at VS or on the ground. When requirements are stated at conditions such as
these, the Levels shall be applied as if the conditions were in the Operational Flight Envelope.
4.1.7 Aircraft Failure States.
The contractor shall define and tabulate all Aircraft Failure States which can affect flying qualities. Aircraft Failure
States consist of Aircraft Normal States modified by one or more malfunctions in aircraft components or systems;
for example, a discrepancy between a selected configuration and an actual configuration. Those malfunctions that
result in center-of-gravity positions outside the center-of-gravity envelope defined in 4.1.1 shall be included. Each
mode of failure shall be considered in all subsequent Flight Phases.
4.1.7.1
Allowable Levels for Aircraft Failure States
4.1.7.2
Aircraft Special Failure States.
Certain components, systems, or combinations thereof may have extremely remote probabilities of failure during
a given flight. The failures may, in turn, be very difficult to predict with any degree of accuracy. Special Failure
States of this type need not be considered in complying with the requirements of this standard if justification for
considering them as Special Failure States is submitted by the contractor and approved by the procuring activity.
4.1.7.3
Probability calculation.
When Aircraft Failure States (4.1.7) exist, a degradation in flying qualities is permitted only if the probability of
encountering a lower Level than specified is sufficiently small. The contractor shall determine, based on the most
accurate available data, the probability of occurrence of each Aircraft Failure State per flight within the
Operational and Service Flight Envelopes. These determinations shall be based on ____________ except that:
a. All aircraft systems are assumed to be operating for the entire flight, unless clearly operative only for a
shorter period
21
MIL-STD-1797A
TABLE II. Aircraft Normal States.
Flight Phase
Weight
Takeoff
TO
Climb
CL
Cruise
CR
Loiter
LO
Descent
ED
Emergency
Deceleration
DE
Approach
PA
Landing
External
Stores
Thrust
Thrust
Vector
Angle
High Lift
Device
D
Emergency
Descent
Wave-off/
Go-Around
C.G.
WO
L
Air-to-Air Combat
CO
Ground Attack
GA
Weapon Delivery/
Launch
WD
Aerial Delivery
AD
Aerial Recovery
AR
Reconnaissance
RC
Refuel Receiver
RR
Refuel Tanker
RT
Terrain Following
TF
Antisubmarine
Search
AS
Close Formation
Flying
FF
Catapult Takeoff
CT
22
Wing
Sweep
Wing
Incidence
Landing
Gear
Speed
Brakes
Bomb
Bay or
Cargo
Doors
Stability
Augmentation
Other
MIL-STD-1797A
b. For these calculations, the length of flight shall be ________ hours
c. Each specific failure is assumed to be present at whichever point in the Flight Envelope being
considered is most critical (in the flying qualities sense).
From these Failure State probabilities and effects, the contractor shall determine the overall probability,
per flight, that one or more flying qualities are degraded to Level 2 because of one or more failures. The
contractor shall also determine the probability that one or more flying qualities are degraded to Level 3.
These probabilities shall be less than the values shown in table III.
TABLE III. Levels for Aircraft Failure States.
PROBABILITY OF
ENCOUNTERING
WITHIN OPERATIONAL
FLIGHT ENVELOPE
Level 2 after failure
< ______ per flight
Level 3 after failure
< ______ per flight
WITHIN SERVICE
FLIGHT ENVELOPE
< ______ per flight
4.1.7.4
Generic failure analysis.
The requirements on the effects of specific types of failures, for example propulsion or flight control
system, shall be met on the basis that the specific type of failure has occurred, regardless of its
probability of occurrence. The requirements of this standard on failure transients shall also be met. The
allowable flying quality Levels for each of the Failure States in 4.1.7 are defined as follows: _______. In
addition, flying qualities in the following specific Failure States shall be as follows:
Failure
Level
___________
_____
4.1.7.5
When Levels are not specified.
Within the Operational and Service Flight Envelopes, all requirements that are not identified with specific
Levels shall be met under all conditions of component and system failure except approved Aircraft
Special Failure States (4.1.7.2).
4.1.7.6
Failures outside the Service Flight Envelopes.
Failures to be considered outside the Service Flight Envelopes but within the corresponding Permissible
Flight Envelopes are _________. After these failures it shall be possible to return safely to the Service
and Operational Flight Envelopes.
4.1.8 Dangerous flight conditions.
Dangerous conditions may exist at which the aircraft should not be flown. When approaching these flight
conditions, it shall be possible by clearly discernible means for the pilot to recognize the impending
dangers and take preventive action. Whenever failures occur that require or limit any flight crew action or
decision concerning flying the aircraft, the crew member concerned shall be given immediate and easily
interpreted indication.
4.1.8.1
Warning and indication.
Warning and indication of approach to a dangerous condition shall be clear and unambiguous. For
example, a pilot must be able to distinguish readily among stall warning (which requires pitching down or
increasing speed), Mach buffet (which may indicate a need to decrease speed), and normal aircraft
vibration (which indicates no need for pilot action).
23
MIL-STD-1797A
4.1.8.2
Devices for indication, warning, prevention, and recovery.
It is intended that dangerous flight conditions be eliminated and the requirements of this standard met by
appropriate aerodynamic design and mass distribution, rather than through incorporation of a special
device or devices. As a minimum, these devices shall perform their function whenever needed but shall
not limit flight within the Operational Flight Envelope. Neither normal nor inadvertent operation of such
devices shall create a hazard to the aircraft. For Levels 1 and 2, nuisance operation shall not be possible.
Functional failure of the devices shall be indicated to the pilot.
4.1.9 Interpretation of subjective requirements.
In several instances throughout the standard, subjective terms such as objectionable flight characteristics,
realistic time delay, normal pilot technique and excessive loss of altitude or buildup of speed, have been
employed where insufficient information exists to establish absolute quantitative criteria. Final
determination of compliance with requirements so worded will be made by the procuring activity.
4.1.10 Interpretation of quantitative requirements.
Many of the numerical requirements of this standard are stated in terms of a linear mathematical
description of classical aircraft. Certain factors, for example flight control system nonlinearities and higherorder dynamics or aerodynamic nonlinearities, can cause an appreciable difference in the aircraft
response apparent to the pilot from that of the linear model of the basic airframe. The contractor shall
determine equivalent classical systems which have responses most closely matching those of the actual
aircraft. Then those numerical requirements of section 4 which are stated in terms of linear system
parameters (such as frequency, damping ratio and modal phase angles) apply to the parameters of that
equivalent system rather than to any particular modes of the higher-order system representation. The
adequacy of the response match between equivalent and actual aircraft, or alternative criteria, shall be
agreed upon by the contractor and the procuring activity. Nonlinearities or higher-order dynamics that
may exist shall not result in any objectionable (for Levels 1 and 2) or dangerous characteristics.
4.1.11 General flying qualities requirements
4.1.11.1 Buffet.
Within the boundaries of the Operational Flight Envelope, there shall be no objectionable buffet which
might detract from the effectiveness of the aircraft in executing its intended missions. In the Permissible
Flight Envelope _____________.
4.1.11.2 Release of store.
The intentional release or ejection of any stores shall not result in objectionable flight characteristics or
impair tactical effectiveness for Levels 1 and 2. However, the intentional release or ejection of stores shall
never result in dangerous or intolerable flight characteristics. This requirement applies for all flight
conditions and store loadings at which normal or emergency release or ejection of the store is
permissible.
4.1.11.3 Effects of armament delivery and special equipment.
Operation of movable parts such as bomb bay doors, cargo doors, armament pods, refueling devices and
rescue equipment, or firing of weapons, release of bombs, or delivery or pickup of cargo shall not cause
buffet, trim changes, or other characteristics which impair the tactical effectiveness of the aircraft under
any pertinent flight condition. These requirements shall be met for Levels 1 and 2.
4.1.11.4 Failures.
No single failure of any component or system shall result in dangerous or intolerable flying qualities;
Special Failure States (4.1.7.2) are excepted. The crew member concerned shall be given immediate and
easily interpreted indications whenever failures occur that require or limit any flight crew action or
decision. The aircraft motions following sudden aircraft system or component failures shall be such that
dangerous conditions can be avoided by the pilot, without requiring unusual or abnormal corrective
action. A realistic time delay of at least __________ between the failure and initiation of pilot corrective
24
MIL-STD-1797A
action shall be incorporated when determining compliance. This time delay shall include an interval
between the occurrence of the failure and the occurrence of a cue such as acceleration, rate,
displacement, or sound that will definitely indicate to the pilot that a failure has occurred, plus an
additional interval which represents the time required for the pilot to diagnose the situation and initiate
corrective action.
Additional requirements apply to transients from propulsion system (4.5.8.4, 4.5.9.5.5, 4.6.5.1, 4.6.6.2,
4.6.7.8) and flight control system (4.2.6.1, 4.2.8.6.5, 4.5.7.1, 4.5.9.5.6. 4.6.5.2, 4.6.7.9) failures.
4.1.11.5 Control margin.
Aerodynamic control power, control surface rate and hinge moment capability shall be sufficient to assure
safety throughout the combined range of all attainable angles of attack (both positive and negative) and
sideslip. This requirement applies to the prevention of loss of control and to recovery from any situation,
including deep stall trim conditions, for all maneuvering, including pertinent effects of factors such as pilot
strength, regions of control-surface-fixed instability, inertial coupling, fuel slosh, the influence of symmetric
and asymmetric stores (4.1.3), stall/post-stall/spin characteristics (4.8.4 through 4.8.4.3.2), atmospheric
disturbances (4.9.1) and Aircraft Failure States (4.1.7 through 4.1.7.6);failure transients and maneuvering
flight appropriate to the Failure State are to be included. Consideration shall be taken of the degree of
effectiveness and certainty of operation of limiters, c.g. control malfunction or mismanagement, and
transients from failures in the propulsion, flight control and other relevant systems.
4.1.11.6 Pilot-induced oscillations (PIO).
There shall be no tendency for pilot-induced oscillations, that is, sustained or uncontrollable oscillations
resulting from the efforts of the pilot to control the aircraft. More specific requirements are in 4.2.1.2,
4.2.2, 4.5.2 and 4.6.3.
4.1.11.7 Residual oscillations.
Any sustained residual oscillations in calm air shall not interfere with the pilot's ability to perform the tasks
required in service use of the aircraft. More specific quantitative requirements are in 4.2.3.
4.1.11.8 Control cross-coupling.
No controller shall create a secondary response which is objectionable (for Levels 1 and 2) or dangerous
(for Level 3). This requirement applies to all continuous and discrete controllers which affect the motion of
the aircraft.
4.1.12 General flight control system characteristics.
As used in this standard, the term flight control system includes the pitch, roll and yaw controls, direct
force controls including leading-edge and trailing-edge flaps, stability augmentation systems, trim
selectors and all mechanisms and devices that they operate, including the feel system. The requirements
of this section, which are directly related to flying qualities, are in addition to the applicable control system
design specification, for example MIL-F-87242 or MIL-C-18244. Some of the important mechanical
characteristics of control systems (including servo valves and actuators) are: friction and preload, lost
motion, flexibility, mass imbalance and inertia, nonlinear gearing, and rate limiting. Meeting separate
requirements on these items, however, will not necessarily ensure that the overall system will be
adequate; the mechanical characteristics must be compatible with the nonmechanical portions of the
control system and with the airframe dynamic characteristics.
4.1.12.1 Control centering and breakout forces.
Pitch, roll, yaw, direct lift and sideforce controls shall exhibit positive centering in flight at any normal trim
setting. Although absolute centering is not required, the combined effects of centering, breakout force,
stability and force gradient shall not produce objectionable flight characteristics, such as poor precisiontracking ability, or permit large departures from trim conditions with controls free. Requirements for
particular controllers are to be found in 4.2.8.5, 4.3.4, 4.5.9.4 and 4.6.7.11.
25
MIL-STD-1797A
4.1.12.2 Cockpit control free play.
The free play in each cockpit control, that is, any motion of the cockpit control which does not move the
control surface in flight, shall not result in objectionable flight characteristics, particularly for smallamplitude control inputs.
4.1.12.3 Adjustable controls.
When a cockpit control is adjustable for pilot physical dimensions or comfort, the control forces defined in
3.4.4 refer to the mean adjustment. A force referred to any other adjustment shall not differ by more than
10 percent from the force referred to the mean adjustment.
4.1.12.4 Rate of control displacement.
The ability of the aircraft to perform the operational maneuvers required of it shall not be limited by control
surface deflection rates in the atmospheric disturbances specified in 4.9.1. Control rates shall be
adequate to retain stabilization and control in the Severe disturbances of those sections. For powered or
boosted controls, the effect of engine speed and the duty cycle of both primary and secondary control
together with the pilot control techniques shall be included when establishing compliance with this
requirement.
4.1.12.5 Dynamic characteristics.
A linear or smoothly varying aircraft response to cockpit-control deflection and to control force shall be
provided for all amplitudes of control input. The response of the control surfaces in flight shall not lag the
cockpit-control force inputs by more than the angles specified, for frequencies equal to or less than the
frequencies specified: ____________.
4.1.12.6 Damping.
All control system oscillations apparent to the pilot shall be well damped, unless they are of such an
amplitude, frequency and phasing that they do not result in objectionable oscillations of the cockpit
controls or the airframe on the ground, during flight and in atmospheric disturbances.
4.1.12.7 Transfer to alternate control modes.
The transient motions and trim changes resulting from the intentional engagement or disengagement of
any portion of the flight control system by the pilot shall be such that dangerous flying qualities never
result. Allowable transients are further specified in 4.2.6.2, 4.2.8.6.6, 4.5.7.2. 4.5.9.5.7, 4.6.5.3, and
4.6.7.10.
4.1.12.8 Flight control system failures.
The following events shall not cause dangerous or intolerable flying qualities:
a. Complete or partial loss of any function of the flight control system as a consequence of any
single failure (approved Aircraft Special Failure States excepted)
b. Failure-induced transient motions and trim changes either immediately upon failure or upon
subsequent transfer to alternate modes
c.
Configuration changes required or recommended following failure.
The crew member concerned shall be provided with immediate and easily interpreted indication whenever
failures occur that require or limit any flight crew action or decision. Allowable transients are specified by
axis in 4.2.6.1, 4.5.7.1 and 4.6.5.2.
4.1.12.9 Augmentation systems.
Operation of stability augmentation and control augmentation systems and devices shall not introduce
any objectionable flight or ground handling characteristics. Any performance degradation of stability and
control augmentation systems due to saturation of components, rate limiting, or surface deflections, shall
be only momentary, and shall not introduce any objectionable flight or ground handling characteristics.
26
MIL-STD-1797A
This requirement particularly applies for all Normal States and Failure states in the atmospheric
disturbances of 4.9.1 and 4.9.2 and during maneuvering flight at the angle-of-attack, sideslip, and loadfactor limits of the Permissible Envelope. It also applies to post-stall gyrations, spins, and recoveries with
all systems, such as the hydraulic and electrical systems, operating in the state that may result from the
gyrations encountered.
4.1.12.10 Auxiliary dive recovery devices.
Operation of any auxiliary device intended solely for dive recovery shall always produce a positive
increment of normal acceleration, but the total normal load factor shall never exceed 0.8nL, controls free.
4.1.12.11 Direct force controllers.
Direct force controllers include direct lift control systems and lateral translation systems. Direct force
controllers which are separate from the attitude controllers shall have a direction of operation consistent
with the sense of the aircraft motion produced, be conveniently and accessibly located, comfortable to
use and compatible with pilot force and motion capabilities. Transients encountered with engagement of
these modes shall meet the requirements of 4.1.12.7, 4.2.6.2, and 4.6.5.3. Functions should be provided
in the control system that would only allow this mode to be engaged within its design flight regime or
maneuvers. When used either by themselves or in combination with other control modes, flight safety and
mission effectiveness shall not be degraded. These systems shall not defeat limiters that are necessary
for stable and controlled flight, or for structural considerations.
4.1.13 General trim requirements
4.1.13.1 Trim system irreversibility.
All trimming devices shall maintain a given setting indefinitely unless changed by the pilot, or by a special
automatic interconnect (such as to the landing flaps), or by the operation of an augmentation device. If an
automatic interconnect or augmentation device is used in conjunction with a trim device, provision shall
be made to ensure the accurate return of the device to its initial trim position on removal of each
interconnect or augmentation command.
4.1.13.2 Rate of trim operation.
Trim devices shall operate rapidly enough to enable the pilot to maintain low control forces under
changing conditions normally encountered in service, yet not so rapidly as to cause oversensitivity or trim
precision difficulties under any conditions, including:
a. Dives and ground attack maneuvers required in normal service operation
b. Level-flight accelerations at maximum augmented thrust from 250 knots or VR/C, whichever is
less, to Vmax at any altitude when the aircraft is trimmed for level flight prior to initiation of the maneuver.
4.1.13.3 Stalling of trim systems.
Stalling of a trim system due to aerodynamic loads during maneuvers shall not result in an unsafe
condition. Specifically, the entire trim system shall be capable of operating during the dive recoveries of
4.2.8.6.3 at any attainable permissible n, at any possible position of the trimming device.
4.1.13.4 Transients and trim changes.
The transients and steady-state trim changes for normal operation of control devices such as throttle,
thrust reversers, flaps, slats, speed brakes, deceleration devices, dive recovery devices, wing sweep and
landing gear shall not impose excessive control forces to maintain the desired heading, altitude, attitude,
rate of climb, speed or load factor without use of the trimmer control. This requirement applies to all inflight configuration changes and combinations of changes made under service conditions, including the
effects of asymmetric operations such as unequal operation of landing gear, speed brakes, slats or flaps.
In no case shall there be any objectionable buffeting or oscillation caused by such devices. More specific
requirements on such control devices are contained in 4.2.5 and 4.1.12.10.
27
MIL-STD-1797A
4.1.13.5 Trim for asymmetric thrust.
For all multi-engine aircraft, it shall be possible to trim the cockpit-control forces to zero in straight flight
with up to two engines inoperative following asymmetric loss of thrust from the most critical propulsive
factors (4.6.5.1). This requirement defines Level 1 in level-flight cruise at speeds from the maximumrange speed for the engine(s)-out configuration to the speed obtainable with normal rated thrust on the
functioning engine(s). Systems completely dependent on the failed engines shall also be considered
failed.
4.1.13.6 Automatic trim system.
Automatic trimming devices shall not degrade or inhibit the action of response limiters.
4.2
Flying qualities requirements for the pitch axis.
Control force and deflection have both been found universally to be important pilot cues, so both controlsfixed and controls-free (control deflection and force) characteristics have been specified. Although control
of the flight path may be the pilot's ultimate aim, pitch attitude control is commonly used as a surrogate.
4.2.1
Pitch attitude dynamic response to pitch controller
4.2.1.1
Long-term pitch response.
Any oscillation with a period of 15 seconds or longer shall have the following damping: ________. Except
as may be provided in 4.4.1, 4.4.1.1 and 4.2.1.2, no aperiodic flight path divergence is allowed within the
Service Flight Envelope for any Level of flying qualities. These requirements apply with cockpit controls
fixed and with them free.
4.2.1.2
Short-term pitch response
The short-term pitch response shall meet the following requirements for control inputs of all magnitudes
that might be experienced in service use: ___________
4.2.2 Pilot-induced pitch oscillations.
The pitch attitude response dynamics of the airframe plus control system shall not change abruptly with
the motion amplitudes of pitch, pitch rate or normal acceleration unless it can be shown that this will not
result in a pilot-induced oscillation. The total phase angle by which normal acceleration measured at the
pilot's location lags the pilot's pitch control force input at a criterion frequency, ωR, shall be less than
_________.
4.2.3 Residual pitch oscillations.
In calm air, any sustained residual oscillations shall not interfere with the pilot's ability to perform the tasks
required in service use of the aircraft. For Levels 1 and 2, oscillations in normal acceleration at the pilot
station greater than __________ will be considered excessive for any Flight Phase. These requirements
shall apply with the pitch control fixed and with it free.
4.2.4 Normal acceleration at pilot station.
Normal acceleration at the pilot station due to pitch control inputs shall have the following characteristics:
____________
4.2.5 Pitch trim changes.
The pitch trim changes caused by operation of other control devices shall not be so large that a peak
pitch control force in excess of 10 pounds for center-stick controllers or 20 pounds for wheel controllers is
required when such configuration changes are made in flight under conditions representative of
operational procedure. Generally, the conditions of table IV will suffice for determination of compliance
with this requirement. With the aircraft trimmed for each specified initial condition, and no retrimming, the
peak force required to maintain the specified parameter constant following the specified configuration
change shall not exceed the stated value for a time interval of at least 5 seconds following the completion
of the pilot action initiating the configuration change. The magnitude and rate of trim change subsequent
28
MIL-STD-1797A
to this time period shall be easily trimmable by use of the normal trimming devices. These requirements
define Level 1. For Levels 2 and 3, the allowable forces are increased by 50 percent ___________.
4.2.6
Pitch axis response to other inputs
4.2.6.1
Pitch axis response to failures, controls free.
With controls free, the aircraft motions due to partial or complete failure of any subsystem of the aircraft
shall not exceed the following limits for at least _________ seconds following the failure: _____________.
4.2.6.2
Pitch axis response to configuration or control mode change.
The transient motions and trim changes resulting from configuration changes or the intentional
engagement or disengagement of any portion of the primary flight control system in equilibrium flight due
to pilot action shall be such that dangerous flying qualities never result. With controls free, the motion
transients resulting from these situations shall not exceed the following limits for at least
___________seconds following the transfer: _______________. These requirements apply only for
Aircraft Normal States (4.1.6).
29
MIL-STD-1797A
TABLE IV. Pitch trim change conditions.
INITIAL TRIM CONDITIONS
FLIGHT
PHASE
ALTITUDE
SPEED
LANDING
GEAR
HIGH-LIFT
DEVICES &
WING FLAPS
30
THRUST
CONFIGURATION
CHANGE
PARAMETER
TO BE HELD
CONSTANT
MIL-STD-1797A
4.2.7
Pitch axis control power
4.2.7.1
Pitch axis control power in unaccelerated flight.
In steady 1g flight at all Service altitudes, the attainment and holding of all speeds between VS and Vmax
shall not be limited by the effectiveness of the pitch control.
4.2.7.2
Pitch axis control power in maneuvering flight.
Within the Operational Flight Envelope it shall be possible to develop, by use of the pitch control alone,
the following range of load factors: _________. This maneuvering capability is required at constant
altitude at the 1g trim speed and, with trim and throttle settings not changed by the Crew, over a range
about the trim speed the lesser of 15 percent or 50 kt equivalent airspeed (except where limited by the
boundaries of the Operational Flight Envelope).
4.2.7.2.1 Load factor response.
The time required to change from one level of normal load factor to another, in pullups and in turning
flight, shall be adequate for all maneuvers appropriate to the Flight Phase, for all conditions within the
Service Flight Envelope. Overshoots that result from abrupt pullups into the lift- or control-system-limited
region of the load factor boundary shall not result in departure or exceedance of load factor limits.
4.2.7.3
Pitch axis control power in takeoff.
The effectiveness of the pitch control shall not restrict the takeoff performance of the aircraft and shall be
sufficient to prevent overrotation during all types of takeoff. It shall be possible to obtain and maintain the
following attitudes during the takeoff roll: _________. For catapult takeoffs __________.
4.2.7.4
Pitch axis control power in landing.
The pitch control shall be sufficiently effective in the landing Flight Phase in close proximity to the ground
that __________________ .
4.2.8
Pitch axis control forces
4.2.8.1
Pitch axis control forces--steady-state control force per g.
These requirements apply for all local gradients throughout the range of Service load factors defined in
4.1.4.2. The term gradient does not include that portion of the force versus n curve within the breakout
force.
a. In steady turning flight and in pullups and pushovers at constant speed, the incremental control
force required to maintain a change in normal load factor and pitch rate shall be in the same sense (aft-more positive, forward--more negative) as that required to initiate the change.
b. The variations in pitch controller force with steady-state normal acceleration shall have no
objectionable nonlinearities within the following load factor range: ________. Outside this range, a
departure from linearity resulting in a local gradient which differs from the average gradient for the
maneuver by more than 50 percent is considered excessive, except that larger increases in force gradient
are permissible at load factors greater than 0.85 nL.
c. The local force gradients shall be: ___________. In addition, FS/n should be near the Level 1
upper boundaries of these gradients for combinations of high frequency and low damping.
For all types of controllers, the control force gradients shall produce suitable flying qualities.
4.2.8.2
Pitch axis control forces--transient control force per g.
The buildup of control forces during maneuver entry must not lag the buildup of normal acceleration at the
pilot's location. In addition, the frequency response of normal acceleration at the pilot station to pitch
control force input shall have the following characteristics: _______________.
31
MIL-STD-1797A
4.2.8.3
Pitch axis control forces--control force variations during rapid speed changes.
When the aircraft is accelerated and decelerated rapidly through the operational speed range and through
the transonic speed range by the most critical combination of changes in power, actuation of deceleration
devices, steep turns and pullups, the magnitude and rate of the associated trim change shall not be so
great as to cause difficulty in maintaining the desired load factor by normal pilot techniques.
4.2.8.4
Pitch axis control forces--control force vs. control deflection.
The gradient of pitch-control force per unit of incremental pitch-control deflection shall be within the
following range: _______. In steady turning flight and in pullups and pushovers at constant speed, the
incremental control position shall be in the same sense (aft--more positive, forward--more negative) as
that required to initiate the change. Dynamically, throughout the frequency range of pilot control inputs the
deflection of the pilot's control must not lead the control force.
4.2.8.5
Pitch axis control breakout forces.
Breakout forces, including friction, preload, etc., shall be within the following limits: __________. These
values refer to the cockpit control force required to start movement of the control surface.
4.2.8.6
Pitch axis control force limits
4.2.8.6.1 Pitch axis control force limits--takeoff.
With the trim setting optional but fixed, the pitch-control forces required during all types of takeoffs for
which the aircraft is designed, including short-field takeoffs and assisted takeoffs such as catapult or
rocket-augmented, shall be within the following limits: ____________. The term takeoff includes the
ground run, rotation, and lift-off, the ensuing acceleration to Vmax(TO), and the transient caused by assist
cessation. Takeoff encompasses operation both in and out of ground effect. Takeoff power shall be
maintained until Vmax(TO) is reached, with the landing gear and high-lift devices retracted in the normal
manner at speeds from Vomin (TO) to Vomax (TO).
4.2.8.6.2 Pitch axis control force limits—landing.
The pitch control forces for landing shall be less than _________ with the aircraft trimmed for the
approach Flight Phase at the recommended approach speed. This limit applies both in and out of ground
effect.
4.2.8.6.3 Pitch axis control force limits--dives.
With the aircraft trimmed for level flight at speeds throughout the Service Flight Envelope, the control
forces in dives to all attainable speeds within the Service Flight Envelope shall not exceed _(a)_. In
similar dives, but with use of trim following the dive entry, it shall be possible with normal piloting
techniques to maintain the forces within: _(b)_.
With the aircraft trimmed for level flight at VMAT but with use of trim optional in the dive, it shall be possible
to maintain the pitch control force within the following limits in dives to all attainable speeds within the
Permissible Flight Envelope_(c)_. The force required for recovery from these dives shall not exceed
___(d)_. Trim and deceleration devices, etc., may be used to assist in recovery if no unusual pilot
technique is required.
4.2.8.6.4 Pitch axis control force limits-sideslips.
With the aircraft trimmed for straight, level flight with zero sideslip, the pitch-control force required to
maintain constant speed in steady sideslips with up to _(a)_ pounds of pedal force in either direction, and
in sideslips as specified in the Operational Flight Envelope, shall not exceed the pitch-control force that
would result in a 1g change in normal acceleration. In no case, however, shall the pitch-control force
exceed_(b)_. If a variation of pitch-control force with sideslip does exist, it is preferred that increasing pull
force accompany increasing sideslip, and that the magnitude and direction of the force change be the
same for right and left sideslips. For Level 3, throughout the Service Flight Envelope there shall be no
uncontrollable pitching motions associated with the sideslip maneuvers discussed above.
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MIL-STD-1797A
4.2.8.6.5 Pitch axis control force limits-failures.
Without retrimming, the change in longitudinal control force required to maintain constant attitude
following complete or partial failure of the flight control system shall not exceed the following limits:
_________.
4.2.8.6.6 Pitch axis control force limits-control mode change.
Without retrimming, the control force changes resulting from intentional engagement or disengagement of
any portion of the primary flight control system by the pilot shall not exceed the following limits:
__________.
4.2.8.7
Pitch axis trim systems.
In straight flight, throughout the Operational Flight Envelope the trimming system shall be capable of
reducing the steady-state control forces to ____________. Pitch trim systems shall not defeat other
features incorporated in the flight control system that prevent or suppress departure from controlled flight
or exceedance of structural limits, or that provide force cues which warn of approach to flight limits. The
failures to be considered in applying Level 2 and 3 requirements shall include trim sticking and runaway in
either direction. It is permissible to meet Level 2 and 3 requirements by providing the pilot with alternate
trim mechanisms or override capability.
4.2.9
Pitch axis control displacements
4.2.9.1
Pitch axis control displacements-takeoff.
With the trim setting optional but fixed, the pitch-control travel during all types of takeoffs for which the
aircraft is designed shall not exceed _________ percent of the total travel, stop-to-stop. Here the term
takeoff includes ground run, rotation and lift-off, the ensuing acceleration to Vmax(TO), and any transient
caused by assist cessation. Takeoff power shall be maintained until Vmax(TO) is reached, with the landing
gear and high-lift devices retracted in the normal manner at speeds from Vomin(TO) to Vmax(TO).
Satisfactory takeoffs, including catapult takeoffs where applicable, shall not depend upon use of the trim
controller during takeoff or upon complicated control manipulation by the pilot.
4.2.9.2
Pitch axis control displacements-maneuvering.
For all types of pitch controllers, the control motions in maneuvering flight shall not be so large or so small
as to be objectionable. In steady turning flight and in pullups at constant speed, the incremental control
deflection required to maintain a change in normal load factor and pitch rate shall be in the same sense
(aft-more positive, forward-more negative) as those required to initiate the change.
4.3
Flying qualities requirements for the normal (flight path) axis
4.3.1
Flight path response to attitude change
4.3.1.1
Transient flight path response to attitude change.
The relation of the flight path response to pitch attitude, for pilot control inputs, shall be as follows:
a. The short-term flight path response to attitude changes shall have the following characteristics:
_______.
b. If a designated controller other than attitude is the primary means of controlling flight path, the
flight path response to an attitude change can be degraded to the following: __________
c. In all cases the pitch attitude response must lead the flight path angle by ___________and must
have a magnitude equal to or greater than the flight path angle.
4.3.1.2
Steady-state flight path response to attitude change.
For flight path control primarily through the pitch attitude controller, the steady-state path and airspeed
response to attitude inputs shall be as follows: ___________. For flight control modes using another
designated flight path control the required flight path response to attitude changes is _________.
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MIL-STD-1797A
4.3.2 Flight path response to designated flight path controller.
When a designated flight path controller (other than the pitch controller) is used as a primary flight path
controller, the short-term flight path response to designated flight path controller inputs shall have the
following characteristics: ________. At all flight conditions the pilot-applied force and deflection required
to maintain a change in flight path shall be in the same sense as those required to initiate the change.
4.3.3
Flight path control power
4.3.3.1
Control power for designated primary flight path controller.
If a separate controller (other than the pitch controller) is provided for primary control of direct lift or flight
path, it shall be capable of producing the following changes in flight path following full actuation of the
controller: _______. This shall be accomplished with pitch attitude held fixed and controls trimmed for an
airspeed of _________.
4.3.3.2
Control power for designated secondary flight path controller.
The secondary controller shall be sufficient to produce the following changes in flight path: ___________.
4.3.4 Flight path controller characteristics.
The breakout, centering, and force gradient characteristics of the designated flight path controller shall be
within the following limits:
Breakout: ± __________lb
Centering: ± _________% of full travel
Force gradient: _________
4.4
Flying qualities requirements for the longitudinal (speed) axis
4.4.1 Speed response to attitude changes.
The correlation between airspeed and pitch attitude shall be as follows: (a)__.
Cockpit controls fixed or free, for Levels 1 and 2 there shall be no tendency for the airspeed to diverge
aperiodically when the aircraft pitch attitude is disturbed from trim. For Level 3, the airspeed divergence
must be within the following limits: ____(b)____.
4.4.1.1
Speed response to attitude changes-relaxation in transonic flight.
The requirements of 4.4.1 may be relaxed in the transonic speed range as follows: __________. This
relaxation does not apply to Level 1 for any Flight Phase which requires prolonged transonic operation.
4.5
Flying qualities requirements for the roll axis
4.5.1
Roll response to roll controller
4.5.1.1
Roll mode.
The equivalent roll mode time constant, TR, shall be no greater than the following: ____________.
4.5.1.2
Spiral stability.
The combined effects of spiral stability, flight control system characteristics and rolling moment change
with speed shall be such that following a disturbance in bank of up to 20 degrees, the time to double
bank-angle amplitude is no less than ______ seconds. This requirement shall be met with the aircraft
trimmed for wings-level, zero-yaw-rate flight and the cockpit controls free.
4.5.1.3
Coupled roll-spiral oscillation.
A coupled roll-spiral mode will be permitted only if it has the following characteristics: __________.
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MIL-STD-1797A
4.5.1.4
Roll oscillations.
With yaw controls free, the response to roll control commands shall meet the following requirements:
________.
4.5.1.5
Roll time delay.
The value of the equivalent time delay, τep, shall be no greater than ___________.
4.5.2 Pilot-induced roll oscillations.
There shall be no tendency for sustained or uncontrollable roll oscillations resulting from efforts of the
pilot to control the aircraft. Specifically, __________.
4.5.3 Linearity of roll response to roll controller.
There shall be no objectionable nonlinearities in the variation of rolling response with roll control
deflection or force. Sensitivity or sluggishness in response to small control deflections or force shall be
avoided.
4.5.4 Lateral acceleration at the pilot station.
With yaw control free and with it used to coordinate turn entry, the ratio of maximum lateral acceleration at
the pilot station to maximum roll rate shall not exceed ___________ for the first 2-1/2 seconds following a
step roll control input.
4.5.5 Roll characteristics in steady sideslip.
In the straight, steady sideslips of 4.6.1.2:
a. An increase in right bank angle, or no change, shall accompany an increase in right sideslip, and
an increase in left bank angle, or no change, shall accompany an increase in left sideslip.
b. Right or zero roll-control deflection and force shall accompany right sideslips, and left or zero rollcontrol deflection and force shall accompany left sideslips. For Levels 1 and 2, the variation of roll- control
deflection and force with sideslip angle shall be essentially linear. This requirement may, if necessary, be
excepted for waveoff (go-around) if task performance is not impaired and no more than 50 percent of rollcontrol power available to the pilot, and no more than 10 pounds of roll-control force, are required in a
direction opposite to that specified herein.
4.5.6 Roll axis control for takeoff and landing in crosswinds.
It shall be possible to take off and land with normal pilot skill and technique in the 90-degree crosswinds
of 4.6.4.
4.5.7
Roll axis response to other inputs
4.5.7.1
Roll axis response to augmentation failures.
With controls free, for at least _________ seconds following the failure the aircraft motions due to partial
or complete failure of the augmentation system shall not exceed the following limits: ___________.
4.5.7.2
Roll axis response to configuration or control mode change.
The transient motions and trim changes resulting from configuration changes or the engagement or
disengagement of any portion of the primary flight control system shall be such that dangerous flying
qualities never result. With controls free, the motion transients resulting from these situations shall not
exceed the following limits for at least ___________ seconds following the transfer: __________. These
requirements apply only for Aircraft Normal States.
4.5.8
Roll axis control power
4.5.8.1
Roll axis response to roll control inputs.
The response to full roll control input shall have the following characteristics: ___________. These
requirements apply throughout the applicable speed-altitude-load-factor flight envelopes, except that the
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MIL-STD-1797A
structural limits on combined rolling and normal acceleration need not be exceeded. For rolls from steady
banked flight, the initial condition shall be coordinated, that is, zero sideslip. The requirements apply to
roll commands to the right and to the left, initiated both from steady bank angles and from wings-level,
straight flight except as otherwise stated.
Inputs are to be abrupt, with time measured from the initiation of control force. The pitch control is to be
held fixed throughout the maneuver. Yaw control pedals shall remain free for Class IV aircraft for Level 1,
and for all carrier-based aircraft in Category C Flight Phases for Levels 1 and 2; but otherwise, yaw
control pedals may be used to reduce sideslip that retards roll rate (not to produce sideslip that augments
roll rate) if such control inputs are simple, easily coordinated with roll control inputs and consistent with
piloting techniques for the aircraft Class and mission.
For Flight Phase TO, the time required to bank may be increased proportional to the ratio of the rolling
moment of inertia at takeoff to the largest rolling moment of inertia at landing, for weights up to the
maximum authorized landing weight.
4.5.8.2
Roll axis control power in steady sideslips.
For Levels 1 and 2, positive effective dihedral (right roll control for right sideslip and left roll control for left
sideslip) shall never be so great that more than ___________ percent of roll control power available to
the pilot is required for sideslips which might be encountered in service deployment.
4.5.8.3
Roll axis control power in crosswinds.
Roll control shall be sufficient to perform the following tasks:
a. On a dry surface it shall be possible to taxi at any angle to a ___________ knot wind
b. Roll control power, in conjunction with other normal means of control, shall be adequate to
maintain a straight path during the takeoff run and the landing rollout in crosswinds up to those specified
in 4.6.4.
c. Roll control power shall be adequate to maintain wings level with up to ___________ deg of
sideslip in the power approach. For Level 1 this shall require not more than __________ percent of the
control power available to the pilot.
d. Following sudden asymmetric loss of thrust from any factor during takeoff, approach, landing and
low-altitude parachute extraction, the aircraft shall be safely controllable in roll in the crosswinds of 4.6.4
from the unfavorable direction.
4.5.8.4
Roll axis control power for asymmetric thrust.
Not more than _________ percent of the roll control available to the pilot shall be needed in meeting the
steady-state and dynamic requirements of 4.6.5.1, 4.1.11.4 and 4.6.6.2 for asymmetric loss of thrust from
any single factor.
4.5.8.5
Roll axis control power in dives and pullouts.
Roll control power shall be adequate to maintain wings level without retrimming, throughout the dives and
pullouts of 4.2.8.6.3.
4.5.8.6
Roll axis control power for asymmetric loading.
When initially trimmed with each asymmetric loading at any speed in the Operational Flight Envelope, and
also for hung stores and all asymmetries encountered in normal operation, it shall be possible to maintain
a straight flight path throughout the Operational Flight Envelope. For Category A Flight Phases, roll
control power with these asymmetric loadings shall be sufficient to hold the wings level at the maximum
load factors of 4.2.7.2 with adequate control margin (4.1.11.5).
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MIL-STD-1797A
4.5.9
Roll axis control forces and displacements
4.5.9.1
Roll control displacements.
For aircraft with wheel controllers, the wheel throw necessary to meet the roll performance requirements
specified in 4.5.8 through 4.5.8.6 shall not exceed __________ degrees in either direction. For aircraft
with stick controllers, no physical interference with the pilot's body, seat, or cockpit structure shall prevent
the pilot from obtaining full control authority.
4.5.9.2
Roll axis control forces to achieve required roll performance.
The roll control force required to obtain the rolling performance specified in 4.5.8.1 shall be neither
greater than ____________ nor less than __________.
4.5.9.3
Roll axis control sensitivity.
The roll control force gradient shall have the following characteristics: ________. In case of conflict
between the requirements of 4.5.9.3 and 4.5.9.2, the requirements of 4.5.9.3 shall govern.
4.5.9.4
Roll axis control centering and breakout forces.
Breakout forces, including friction, preload, etc., shall be within the following limits: __________.
4.5.9.5
Roll axis control force limits
4.5.9.5.1 Roll axis control force limits in steady turns.
For Levels 1 and 2, with the aircraft trimmed for wings-level, straight flight it shall be possible to maintain
steady turns in either direction with the yaw controls free at the following combinations of bank angle and
roll controller force characteristics: ____________.
4.5.9.5.2 Roll axis control force limits in dives and pullouts.
Roll control forces shall not exceed __________ lb in dives and pullouts to the maximum speeds
specified in the Service Flight Envelope.
4.5.9.5.3 Roll axis control force limits in crosswinds.
It shall be possible to take off and land in the crosswinds specified in 4.6.4 without exceeding the
following roll control forces: ________
4.5.9.5.4 Roll axis control force limits in steady sideslips.
For Levels 1 and 2, positive effective dihedral (right roll control for right sideslip and left roll control for left
sideslip) shall never be so great that more than ________ pounds of roll-stick force or _________ pounds
of roll-wheel force is required for sideslip angles that might be experienced in service employment. In final
approach the roll control forces shall not exceed ________ lb when in a straight, steady sideslip of
_________ deg.
4.5.9.5.5 Roll axis control force limits for asymmetric thrust.
To meet the steady-state requirements of 4.6.5.1, 4.6.6.2 and 4.1.11.4 shall not require any roll control
force greater than __________.
4.5.9.5.6 Roll axis control force limits for failures.
The change in roll control force required to maintain constant attitude following a failure in the flight
control system shall not exceed ________ pounds for at least five seconds following the failure.
4.5.9.5.7 Roll axis control force limits for configuration or control mode change.
The control force changes resulting from configuration changes or the intentional engagement or
disengagement of any portion of the flight control system shall not exceed the following limits:
_________.
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MIL-STD-1797A
4.6
Flying qualities requirements for the yaw axis
4.6.1
Yaw axis response to yaw and side-force controllers
4.6.1.1
Dynamic lateral-directional response.
The equivalent parameters describing the oscillatory response in sideslip, yaw and roll to a yaw control
input shall have the following characteristics: ________. These requirements shall be met in trimmed
flight with cockpit controls fixed and with them free, and in steady maneuvers, in oscillations of any
magnitude that might be experienced in operational use. If the oscillation is nonlinear with amplitude, the
requirement shall apply to each cycle of the oscillation.
4.6.1.2
Steady sideslips.
The long-term response to yaw-control-pedal deflections shall have the following characteristics:
___________. This requirement applies to yaw-and-roll-control-induced steady, upright, zero-yaw-rate
sideslips with the aircraft trimmed for wings-level straight flight, at sideslip angles up to those produced or
limited by:
a. Full yaw-control-pedal deflection, or
b.
250 pounds of yaw-control-pedal force, or
c.
Maximum roll control or surface deflection,
except that for single-propeller-driven aircraft during waveoff (go-around), yaw-control-pedal deflection in
the direction opposite to that required for wings-level straight flight need not be considered beyond the
deflection for a 10 deg change in sideslip from the wings-level straight flight condition.
Right yaw-control-pedal force and deflection shall produce left sideslips, and left yaw-control-pedal force
and deflection shall produce right sideslips. For Levels 1 and 2 the following requirements shall apply.
The variation of sideslip angle with yaw-control-pedal deflection shall be essentially linear for sideslip
angles between _________ and __________ degrees. The variation of sideslip angle with yaw-controlpedal force shall be essentially linear for sideslip angles between __________ degrees and ___________
degrees. For larger sideslip angles, an increase in yaw-control-pedal force shall always be required for an
increase in sideslip.
4.6.1.3
Wings-level turn.
For a wings-level-turn mode of control, the following requirements apply:
a. Dynamic response to direct force control (DFC) input: The bandwidth of the open-loop response
of heading or lateral flight path angle to the DFC input shall be greater than __________ for Flight Phase
_____________. Turns shall occur at approximately zero sideslip angle and zero or constant small bank
angle when using the DFC controller.
b. Steady-state response to direct force control input: Maximum DFC inputs shall produce at least
_____________
c. Direct force control forces and deflections: Use of the primary DFC shall not require use of
another control manipulator to meet the above dynamic response requirement. The controller
characteristics shall meet the following requirements: _____________
d. Crew acceleration: Abrupt, large DFC inputs shall not produce head or arm motions which
interfere with task performance. Crew restraints shall not obstruct the crew's normal field of view nor
interfere with manipulation of any cockpit control required for task performance.
4.6.2 Yaw axis response to roll controller.
The sideslip excursions to step roll control inputs shall meet the following criteria: ________. Yaw controls
shall be free and, in initial steady turns, fixed at the deflection for zero sideslip in the turn.
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MIL-STD-1797A
4.6.3 Pilot-induced yaw oscillations.
There shall be no tendency for sustained or uncontrollable yaw oscillations resulting from efforts of the
pilot to control the aircraft. Specifically, ____________.
4.6.4 Yaw axis control for takeoff and landing in crosswinds.
It shall be possible to take off and land with normal pilot skill and technique in 90 deg crosswinds from
either side of velocities up to ___________.
4.6.5
Yaw axis response to other inputs
4.6.5.1
Yaw axis response to asymmetric thrust.
Asymmetric loss of thrust may be caused by many factors including engine failure, inlet unstart, propeller
failure or propeller-drive failure. The requirements apply for the appropriate Flight Phases when any
single failure or malperformance of the propulsive system, including inlet, exhaust, engines, propellers, or
drives causes loss of thrust on one or more engines or propellers, considering also the effect of the failure
or malperformance on all subsystems powered or driven by the failed propulsive system. It shall be
possible for the pilot to maintain directional control of the aircraft following a loss of thrust from the most
critical propulsive source, allowing a realistic time delay of __________ seconds, as follows:
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MIL-STD-1797A
Takeoff run:
During the takeoff run it shall be possible to maintain a straight path on the
takeoff surface without deviations of more than _____ feet from the path
originally intended, following sudden asymmetric loss of thrust. For the
continued takeoff, the requirement shall be met when thrust is lost at speeds
from the refusal speed (based on the shortest runway from which the aircraft is
designed to operate) to the maximum takeoff speed, with takeoff thrust
maintained on the operative engine(s); without depending upon release of the
pitch, roll, yaw or throttle controls; and using only controls not dependent upon
friction against the takeoff surface. For the aborted takeoff, the requirement
shall be met at all speeds below the maximum takeoff speed; however,
additional controls such as nose wheel steering and differential braking may be
used. Automatic devices that normally operate in the event of a thrust failure
may be used in either case.
Airborne:
After lift-off, it shall be possible without a change in selected configuration to
achieve straight flight following critical sudden asymmetric loss of thrust at
speeds from Vmin(TO) to Vmax(TO), and thereafter to maintain straight flight
throughout the climbout and to perform 20-degree-banked turns with and
against the inoperative propulsive unit. Automatic devices that normally
operate in the event of a thrust failure may be used, and for straight flight the
aircraft may be banked up to 5 degrees away from the inoperative engine.
Waveoff/go-around:
At any airspeed down to Vmin(PA) it shall be possible to achieve and maintain
steady, straight flight with waveoff (go-around) thrust on the remaining engines
following sudden asymmetric loss of thrust from the most critical factor.
Configuration changes within the capability of the crew while retaining control
of the aircraft, and automatic devices that normally operate in the event of a
propulsion failure, may be used.
Crosswinds:
The aircraft response requirements for asymmetric thrust in takeoff and landing
apply in the crosswinds of 4.6.4 from the adverse direction.
General:
The static directional stability shall be such that at all speeds above with the
critical asymmetric loss of thrust while the other engine(s) develop(s) normal
rated thrust, the aircraft with yaw control pedals free may be balanced
directionally in steady, straight flight. The trim settings shall be those required
for wings-level, straight flight prior to the failure.
4.6.5.2
Yaw axis response to failures.
With controls free, the yawing motions due to failures shall not exceed ________ for at least __________
seconds following the failure.
4.6.5.3
Yaw axis response to configuration or control mode change.
With controls free, the transients resulting from the configuration changes or engagement or
disengagement of any portion of the flight control system due to pilot action shall not exceed _________
for at least ___________ seconds following the transfer. This requirement applies only for Aircraft Normal
States, within the Service Flight Envelope.
4.6.6 Yaw axis control power.
Directional stability and control characteristics shall enable the pilot to balance yawing moments and
control yaw and sideslip.
4.6.6.1
Yaw axis control power for takeoff, landing, and taxi.
The following requirements shall be met:
a. It shall be possible to taxi on a dry surface at any angle to a ________ knot wind.
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MIL-STD-1797A
b. In the takeoff run, landing rollout, and taxi, yaw control power in conjunction with other normal
means of control shall be adequate to maintain a straight path on the ground or other landing surface.
This applies to calm air and in crosswinds up to the values specified in 4.5.6, on wet runways, and on
_____________.For very slippery runways, the requirement need not apply for crosswind components at
which the force tending to blow the aircraft off the runway exceeds the opposing tire-runway frictional
force with the tires supporting all of the aircraft's weight.
c. If compliance with (b) is not demonstrated by test under the adverse runway conditions of (b),
directional control shall be maintained by use of aerodynamic controls alone at all airspeeds above
____________ kt.
d. Yaw axis control power shall be adequate to develop _________ degrees of sideslip in the power
approach.
e. All carrier-based aircraft shall be capable of maintaining a straight path on the ground without the
use of wheel brakes, at airspeeds of 30 knots and above, during takeoffs and landings in a 90-degree
crosswind of at least 0.1 VS(L).
4.6.6.2
Yaw axis control power for asymmetric thrust.
Yaw control power shall be sufficient to meet the requirements of 4.6.5.1. In addition, at the one-engineout speed for maximum range with any engine initially failed, upon failure of the most critical remaining
engine the yaw control power shall be adequate to stop the transient motion and thereafter to maintain
straight flight from that speed to the speed for maximum range with both engines failed. Further, it shall
be possible to effect a safe recovery at any service speed above Vomin(CL) following sudden simultaneous
failure of the two critical engines.
4.6.6.3
Yaw axis control power with asymmetric loading.
Yaw control power shall be sufficient to meet 4.5.8.6.
4.6.7 Yaw axis control forces.
Sensitivity to yaw control pedal forces shall be sufficiently high that directional control and force
requirements can be met and satisfactory coordination can be achieved without unduly high control
forces, yet sufficiently low that occasional improperly coordinated control inputs will not cause a
degradation in flying qualities Level.
4.6.7.1
Yaw axis control force limits in rolling maneuvers.
In the maneuvers described in 4.5.8.1, directional-control effectiveness shall be adequate to maintain
zero sideslip with pedal force not greater than _________ lb.
4.6.7.2
Yaw axis control force limits in steady turns.
It shall be possible to maintain steady coordinated turns in either direction, using _______ deg of bank
with a pedal force not exceeding ________ lb, with the aircraft trimmed for wings-level, straight flight.
These requirements constitute Levels 1 and 2.
4.6.7.3
Yaw axis control force limits during speed changes.
When initially trimmed directionally with symmetric power, the trim change with speed shall be such that
wings-level, straight flight can be maintained over a speed range of ±30 percent of the trim speed or ±100
kt equivalent airspeed, whichever is less (except where limited by boundaries of the Service Flight
Envelope) with yaw-control-pedal forces not greater than _______ lb without retrimming.
4.6.7.4
Yaw axis control force limits in crosswinds.
It shall be possible to take off and land in the crosswinds specified in 4.6.4 without exceeding the
following yaw control forces: ______.
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MIL-STD-1797A
4.6.7.5
Yaw axis control force limits with asymmetric loading.
When initially trimmed directionally with each asymmetric loading specified in 4.1.3 at any speed in the
Operational Flight Envelope, it shall be possible to maintain a straight flight path throughout the
Operational Flight Envelope with yaw-control-pedal forces not greater than _________ lb without
retrimming.
4.6.7.6
Yaw axis control force limits in dives and pullouts.
Throughout the dives and pullouts of 4.2.8.6.3, yaw-control-pedal forces shall not exceed _______ lb in
dives and pullouts to the maximum speeds in the Service Flight Envelope.
4.6.7.7
Yaw axis control force limits for waveoff (go-around).
The response to thrust, configuration and airspeed change shall be such that the pilot can maintain
straight flight during waveoff (go-around) initiated at speeds down to VS(PA) with yaw control pedal forces
not exceeding ________ lb when trimmed to Vomin(PA). The preceding requirements apply for Levels 1
and 2. The Level 3 requirement is to maintain straight flight in these conditions with yaw control pedal
forces not exceeding _______ lb. Bank angles up to 5 degrees are permitted for all Levels.
4.6.7.8
Yaw axis control force limits for asymmetric thrust during takeoff.
The following requirements shall be met:
Takeoff run:
During the takeoff run, to stay within the allowable path deviation of 4.6.5.1, yawcontrol forces shall not exceed _______ lb.
Airborne:
For the continued takeoff, to achieve straight flight following sudden asymmetric loss
of thrust and then maintain straight flight throughout the climbout, as in 4.6.5.1, shall
not require a yaw control pedal force greater than _____ lb.
4.6.7.9
Yaw axis control force limits with flight control failures.
The change in yaw control force required to maintain constant attitude following a failure in the flight
control system shall not exceed _________ lb for at least 5 seconds following the failure.
4.6.7.10 Yaw axis control force limits-control mode change.
The change in yaw control force required to maintain zero sideslip following configuration changes or
engagement or disengagement of any portion of the primary flight control system due to pilot action in
stabilized flight shall not exceed the following limits: ___________. These requirements apply only for
Aircraft Normal States.
4.6.7.11 Yaw axis breakout forces.
Yaw-control breakout forces, including friction, preload, etc., shall be within the following limits: ________.
These values refer to the cockpit control force required to start movement of the surface.
4.7
Flying qualities requirements for the lateral flight path axis
4.7.1 Dynamic response for lateral translation.
The following requirements shall be met:
a. Dynamic response to direct force control input: The bandwidth of the open-loop response of
lateral position to lateral translation control input shall be greater than ________ for Flight Phase
__________. Lateral translations shall occur at essentially zero bank angle and zero change in heading.
b. Steady-state response to lateral translation control input: Maximum control force input shall produce at least ________ degrees of sideslip.
c. Lateral translation control forces and deflections: Use of the primary lateral translation control
shall not require use of another control manipulator to meet requirement a. The controller characteristics
shall meet the following requirements: _____________.
d. Crew accelerations: Abrupt, large control inputs shall not produce head or arm motions which
interfere with task performance. Crew restraints shall not obstruct the crew's normal field of view nor
interfere with manipulation of any cockpit control required for task performance.
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MIL-STD-1797A
4.8
Flying qualities requirements for combined axes
4.8.1 Cross-axis coupling in roll maneuvers.
In yaw-control-free maximum-performance rolls through ____________ degrees, entered from straight
flight or during turns, pushovers, or pullups ranging from 0 g to 0.8 nL, including simultaneous pitch and
roll commands, none of the resulting yaw or pitch motions, or sideslip or angle of attack changes, shall
exceed structural limits or cause other dangerous flight conditions such as uncontrollable motions or roll
autorotation. Rudder pedal inputs used to roll the aircraft with lateral control fixed, or when used in a
coordinated manner with lateral control inputs, shall not result in departures in pitch, roll, or yaw.
During combat-type maneuvers involving rolls through angles up to __________ degrees and rolls which
are checked at any given bank angle up to that value, the yawing and pitching shall not be so severe as
to impair the tactical effectiveness of the maneuver. These requirements define Level 1 and 2 operation.
4.8.2 Crosstalk between pitch and roll controllers.
The pitch- and roll- control force and displacement sensitivities and breakout forces shall be compatible
so that intentional inputs to one control axis will not cause objectionable inputs to the other.
4.8.3 Control harmony.
The following control forces are considered to be limiting values compatible with the pilot's capability to
apply simultaneous forces: _________. Larger simultaneous control forces shall not be required to
perform any customary and expected maneuvers.
4.8.4 Flight at high angle of attack.
4.8.4 through 4.8.4.3.2 concern stall warning, stalls, departures from controlled flight, post-stall gyrations,
spins, recoveries, and related characteristics. They apply at speeds and angles of attack which in general
are outside the Service Flight Envelope. They are intended to assure safety and the absence of mission
limitations due to high-angle-of-attack flight characteristics.
4.8.4.1
Warning cues.
Warning or indication of approach to stall or loss of aircraft control shall be clear and unambiguous.
4.8.4.2
Stalls.
Stall is defined according to 3.4.2 (VS) and 3.4.5 (αS). The stall requirements apply for all Aircraft Normal
States in straight unaccelerated flight and in turns and pullups with attainable normal accelerations up to
nL. Specifically to be evaluated are: __________. Also, the requirements apply to Aircraft Failure States
that affect stall characteristics.
4.8.4.2.1 Stall approach.
The aircraft shall exhibit the following characteristics in the stall approach:
a. The onset of warning of stall approach (4.8.4.1) shall occur within the following speed range for 1g stalls: _________, and within the following range (or percentage) of lift for accelerated stalls:
__________, but not within the Operational Flight Envelope.
b. An increase in intensity of the warning with further increase in angle of attack shall be sufficiently
marked to be noted by the pilot. The warning shall continue until the angle of attack is reduced to a value
less than that for warning onset. Prior to the stall, uncommanded oscillations shall not ___________.
c. At all angles of attack up to the stall, the cockpit controls shall remain effective in their normal
sense, and small control inputs shall not result in departure from controlled flight.
d.
Stall warning shall be easily perceptible and shall consist of ____________.
4.8.4.2.2 Stall characteristics.
The following requirements apply for all stalls, including stalls entered abruptly:
43
MIL-STD-1797A
a. In the unaccelerated stalls of 4.8.4.2, the aircraft shall not exhibit rolling, yawing or downward
pitching at the stall which cannot be controlled to stay within ________ deg.
b. It is desired that no pitch-up tendencies occur in stalls, unaccelerated or accelerated. However,
in the unaccelerated stalls of 4.8.4.2 mild nose-up pitch may be acceptable if no pitch control force
reversal occurs and no dangerous, unrecoverable or objectionable flight conditions result. In the
accelerated stalls of 4.8.4.2, a mild nose-up tendency may be acceptable if the operational effectiveness
of the aircraft is not compromised and the aircraft has adequate stall warning, pitch control effectiveness
is such that it is possible to stop the pitch-up promptly and reduce the angle of attack, and at no point
during the stall approach or recovery does any portion of the aircraft exceed structural limit loads.
4.8.4.2.3 Stall prevention and recovery.
The following requirements shall be met:
a. It shall be possible to prevent the stall by moderate use of the pitch control alone at the onset of
the stall warning.
b. It shall be possible to recover from a stall by simple use of the pitch, roll and yaw controls with
cockpit control forces not to exceed __________, and to regain level flight without excessive loss of
altitude or buildup of speed. Throttles shall remain fixed until an angle of attack below the stall has been
regained, unless compliance would result in exceeding engine operating limitations.
c. In the straight-flight stalls of 4.8.4.2, with the aircraft trimmed at an airspeed not greater than 1.4
VS, pitch control power shall be sufficient to recover from any attainable angle of attack.
d. Operation of automatic departure/spin prevention or recovery devices and flight control modes
shall not interfere with the pilot's ability to prevent or recover from stalls.
4.8.4.2.4 One-engine-out stalls.
On multi-engine aircraft it shall be possible to recover safely from stalls with the critical engine
inoperative. Thrust on the remaining engines shall be at ________.
4.8.4.3
Post-stall gyrations and spins.
The post-stall gyration and spin requirements apply to all modes of motion that can be entered from
upsets, decelerations, and extreme maneuvers appropriate to the Class and Flight Phase Category.
Entries from inverted flight and tactical entries ________ be included. Entry angles of attack and sideslip
up to maximum control capability and under dynamic flight conditions are to be included, except as limited
by structural considerations. Thrust settings up to and including MAT shall be included, with and without
one critical engine inoperative at entry. The requirements hold for all Aircraft Normal States and for all
states of stability and control augmentation systems except approved Special Failure States. Store
release shall not be allowed during loss of control, spin or gyration, recovery, or subsequent dive pullout.
Automatic disengagement or mode-switching of augmentation systems and automatic flight control
system modes, however, is permissible if it is necessary; re-engagement in the normal mode shall be
possible in flight following recovery. Specific flight conditions to be evaluated are: ___________
4.8.4.3.1 Departure from controlled flight.
The aircraft shall be _________ resistant to departure from controlled flight, post-stall gyrations and
spins. Adequate warning of approach to departure (4.8.4.1) shall be provided. The aircraft shall exhibit no
uncommanded motion which cannot be arrested promptly by simple application of pilot control. At all
angles of attack within the Operational Flight Envelope, following sudden asymmetric loss of thrust from
the most critical factor, it shall be possible to avoid departure without exercise of exceptional pilot skill.
4.8.4.3.2 Recovery from post-stall gyrations and spins.
The post-stall characteristics shall be determined. For all aircraft:
a. The proper recovery technique(s) must be readily ascertainable by the pilot and simple and easy
to apply under the motions encountered.
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MIL-STD-1797A
b. A single technique shall provide prompt recovery from all post-stall gyrations and incipient spins.
The same technique, or a compatible one, is required for spin recovery. For all modes of spin that can
occur, these recoveries shall be attainable within ________. Avoidance of a spin reversal or an adverse
mode change shall not depend upon precise pilot control timing or deflection.
c. Operation of automatic stall/departure/spin prevention devices and flight control modes shall not
interfere with or prevent successful recovery of the aircraft by the pilot.
d. Safe and consistent recovery and pullouts shall be accomplished without exceeding the following
forces: _________, and without exceeding structural limitations.
4.9
Flying qualities requirements in atmospheric disturbances
4.9.1 Allowable flying qualities degradations in atmospheric disturbances.
Levels and qualitative degrees of suitability of flying qualities as indicated in 3.3 are used to tailor the
requirements to abnormal conditions such as flight outside the Operational Flight Envelope and Aircraft
Failure States, as well as normal conditions. Abnormalities may also occur in the form of large
atmospheric disturbances, or some combination of conditions. For these factors a degradation of flying
qualities is permitted as specified herein:
a. In atmospheric disturbances the minimum required flying qualities for Aircraft Normal States are
as specified in table V.
b. In atmospheric disturbances the minimum required flying qualities for Aircraft Failure States are
as specified in table VI.
TABLE V. Flying qualities in atmospheric disturbances for Aircraft Normal States.
ATMOSPHERIC
DISTURBANCES
WITHIN OPERATIONAL
FLIGHT ENVELOPE
WITHIN SERVICE
FLIGHT ENVELOPE
LIGHT TO CALM
_______________
_______________
MODERATE TO LIGHT
_______________
_______________
SEVERE TO
MODERATE
_______________
_______________
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MIL-STD-1797A
TABLE VI. Flying qualities in atmospheric disturbances for Aircraft Failure States.
ATMOSPHERIC
DISTURBANCES
FAILURE STATE I*
FAILURE STATE II*
_______________
_______________
_______________
_______________
_______________
_______________
LIGHT TO CALM
MODERATE TO LIGHT
SEVERE TO
MODERATE
* Failure State I: ___________
** Failure State II: ___________
For this purpose atmospheric disturbances are defined separately for high (above approximately 1750 ft)
and low altitudes: _________.
Crosswind intensities at touchdown are defined as: ________.
Required wind-shear capability is: __________.
4.9.2 Definition of atmospheric disturbance model form.
When compliance is to be shown using piloted simulation, an atmospheric disturbance model appropriate
to the piloting task shall be included. As a minimum, the atmospheric disturbance model shall consist of
_________.
4.9.3 Application of disturbance models in analyses.
The gust and turbulence velocities shall be applied to the aircraft equations of motion through the
aerodynamic terms only, and the direct effect on the aerodynamic sensors shall be included when such
sensors are part of the aircraft augmentation system. Application of the disturbance model depends on
the motion variables and the range of frequencies of concern in the analysis. When structural modes are
within or close to this range, the exact distribution of turbulence velocities must be considered. For this
purpose it is acceptable to consider ug and vg as being one-dimensional for the evaluation of aerodynamic
forces and moments. The wg should be considered two-dimensional, a function of both x and y.
When structural modes are not significant to the analysis or simulation, airframe rigid-body responses
may be evaluated by considering uniform gust or turbulence immersion along with linear gradients of the
disturbance velocities. The uniform immersion is accounted for by ug, vg, wg defined at the aircraft center
of gravity. The angular velocities due to turbulence are equivalent in effect to aircraft angular velocities.
Approximations for these angular velocities are defined (precise only at very low frequencies) as follows:
− α& g = q g =
∂w g
∂x
,
pg =
−∂ w g
∂y
,
rg =
−∂ v g
∂x
For altitudes below 1750 ft, the turbulence velocity components ug, vg, and wg are to be taken along axes
corresponding to ug aligned along the horizontal relative mean wind vector and wg vertical.
46
MIL-STD-1797A
5.
VERIFICATION
5.1
General requirements-verification
5.1.1 Loadings-verification.
The contractor shall furnish the required loading data in accordance with the Contract Data Requirements
List.
5.1.2 Moments and products of inertia-verification.
The contractor shall furnish moments and products of inertia data in accordance with the Contract Data
Requirements List.
5.1.3 Internal and external stores-verification.
The analysis, simulation, and testing to verify compliance with this standard shall include the listed stores.
The contractor shall furnish a list of store restrictions in accordance with the Contract Data Requirements
List.
5.1.4
Flight Envelopes-verification
5.1.4.1
Operational Flight Envelopes-verification.
The contractor shall submit the Operational Flight Envelopes for approval by the procuring activity in
accordance with the Contract Data Requirements List.
5.1.4.2
Service Flight Envelopes-verification.
The contractor shall submit the required data for approval by the procuring activity in accordance with the
Contract Data Requirements List.
5.1.4.3
Permissible Flight Envelopes-verification.
The contractor shall provide the required data for approval of the procuring activity in accordance with the
Contract Data Requirements List.
5.1.5 Configurations and States of the aircraft-verification.
The contractor shall furnish a list of aircraft configurations in accordance with the Contract Data
Requirements List.
5.1.6 Aircraft Normal States-verification.
The contractor shall furnish a list of Aircraft Normal States in accordance with the Contract Data
Requirements List.
5.1.6.1
Allowable Levels for Aircraft Normal States-verification.
Verification shall be by analysis, simulation and test.
5.1.6.2
Flight outside the Service Flight Envelopes-verification.
Verification shall be by analysis, simulation and test.
5.1.6.3
Ground operation-verification.
Verification shall be by analysis, simulation and test.
5.1.7 Aircraft Failure States-verification.
The contractor shall furnish the required data in accordance with the Contract Data Requirements List.
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MIL-STD-1797A
5.1.7.1
Allowable Levels for Aircraft Failure States-verification
5.1.7.2
Aircraft Special Failure States-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirement List, for
review by the procuring activity.
5.1.7.3
Probability calculation-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirements List.
5.1.7.4
Generic failure analysis-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirements List.
5.1.7.5
When Levels are not specified-verification.
Verification shall be by analysis, simulation and flight test.
5.1.7.6
Failures outside the Service Flight Envelopes-verification.
The contractor shall review the list furnished by the procuring activity and through its own analysis modify
and extend that list as necessary for adequate coverage of flying qualities degradations, subject to
procuring activity approval, in accordance with the Contract Data Requirements List. Verification of safe
return capability shall be by analysis, simulation, and flight test.
5.1.8 Dangerous flight conditions-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
5.1.8.1
Warning and indication-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
5.1.8.2
Devices for indication, warning, prevention, and recovery-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
5.1.9 Interpretation of subjective requirements-verification.
Verification shall be by analysis, simulation and test.
5.1.10 Interpretation of quantitative requirements-verification.
Verification shall be by analysis, simulation and test.
5.1.11 General flying qualities requirements-verification
5.1.11.1 Buffet-verification.
Verification shall be by analysis, simulation and test.
5.1.11.2 Release of stores-verification.
Verification shall be by analysis, simulation and test.
5.1.11.3 Effects or armament delivery and special equipment-verification.
Verification shall be by analysis, simulation and flight test.
5.1.11.4 Failures-verification.
Verification shall be by analysis, simulation and test.
5.1.11.5 Control margin-verification.
Verification shall be by analysis, simulation, and ground and flight test.
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MIL-STD-1797A
5.1.11.6 Pilot-induced oscillations(PIO)-verification.
Verification shall be by analysis, simulation and flight test.
5.1.11.7 Residual oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.1.11.8 Control cross-coupling-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12 General flight control system characteristics-verification.
Verification shall be by analysis and flight test.
5.1.12.1 Control centering and breakout forces-verification.
Verification shall be by analysis and test. Measurement of breakout forces on the ground will ordinarily
suffice in lieu of actual flight measurement, provided that qualitative agreement between ground
measurement and flight measurement can be established.
5.1.12.2 Cockpit control free play-verification.
Verification shall be by analysis and flight test.
5.1.12.3 Adjustable controls-verification.
Verification shall be by inspection.
5.1.12.4 Rate of control displacement-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.5 Dynamic characteristics-verification.
Verification shall be by analysis and test.
5.1.12.6 Damping-verification.
Verification shall be by analysis and flight test.
5.1.12.7 Transfer to alternate control modes-verification.
Verification shall be by analysis and flight test.
5.1.12.8 Flight control system failures-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.9 Augmentation systems-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.10 Auxiliary dive recovery devices-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.11 Direct force controllers-verification.
Verification shall be by analysis, simulation, inspection, and flight test.
5.1.13 General trim requirements-verification
5.1.13.1 Trim system irreversibility-verification.
Verification shall be by analysis and test.
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MIL-STD-1797A
5.1.13.2 Rate of trim operation-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.3 Stalling of trim systems-verification.
Verification shall be by analysis and flight test.
5.1.13.4 Transients and trim changes-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.5 Trim for asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.6 Automatic trim system-verification.
Verification shall be by analysis, simulation and flight test.
5.2
Flying qualities requirements for the pitch axis-verification
5.2.1
Pitch attitude dynamic response to pitch controller-verification
5.2.1.1
Long-term pitch response-verification.
Verification shall be by analysis, simulation and flight test.
5.2.1.2
Short-term pitch response-verification.
Verification shall be by analysis, simulation and flight test.
5.2.2 Pilot-induced pitch oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.2.3 Residual pitch oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.2.4 Normal acceleration at pilot station-verification.
Verification shall be by analysis, simulation and flight test.
5.2.5 Pitch trim changes-verification.
Verification shall be by analysis, simulation and flight test.
5.2.6
Pitch axis response to other inputs-verification
5.2.6.1
Pitch axis response to failures, controls free-verification.
Verification shall be by analysis, simulation and flight test.
5.2.6.2
Pitch axis response to configuration or control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.2.7
Pitch axis control power-verification
5.2.7.1
Pitch axis control power in unaccelerated flight-verification.
Verification shall be by analysis, simulation and flight test.
5.2.7.2
Pitch axis control power in maneuvering flight-verification.
Verification shall be by analysis, simulation and flight test.
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MIL-STD-1797A
5.2.7.2.1 Load factor response-verification.
Verification shall be by analysis, simulation and flight test.
5.2.7.3
Pitch axis control power in takeoff-verification.
Verification shall be by analysis, simulation and flight test.
5.2.7.4
Pitch axis control power in landing-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8
Pitch axis control forces-verification
5.2.8.1
Pitch axis control forces-steady-state control force per g-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.2
Pitch axis control forces-transient control force per g-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.3
Pitch axis control forces-control force variations during rapid speed changesverification.
Verification shall be by analysis, simulation and flight test.
5.2.8.4
Pitch axis control forces-control force vs. control deflection-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.5
Pitch axis control breakout forces-verification.
Verification shall be by analysis, simulation and test.
5.2.8.6
Pitch axis control force limits-verification
5.2.8.6.1 Pitch axis control force limits-takeoff-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.7
Pitch axis control force limits-landing-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.7.1 Pitch axis control force limits-dives-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.7.2 Pitch axis control force limits-sideslips-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.7.3 Pitch axis control force limits-failures-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.7.4 Pitch axis control force limits-control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.2.8.8
Pitch axis trim systems-verification.
Verification shall be by analysis, simulation and flight test.
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MIL-STD-1797A
5.2.9
Pitch axis control displacements-verification
5.2.9.1
Pitch axis control displacements-takeoff-verification.
Verification shall be by analysis, simulation and flight test.
5.2.9.2
Pitch axis control displacements-maneuvering-verification.
Verification shall be by analysis, simulation and flight test.
5.3
Flying qualities requirements for the normal (flight path) axis-verification
5.3.1
Flight path response to attitude change-verification
5.3.1.1
Transient flight path response to attitude change-verification.
Verification shall be by analysis, simulation and flight test.
5.3.1.2
Steady-state flight path response to attitude change-verification.
Verification shall be by analysis, simulation and flight test.
5.3.2 Flight path response to designated flight path controller-verification.
Verification shall be by analysis, simulation and flight test.
5.3.3
Flight path control power-verification
5.3.3.1
Control power for designated primary flight path controller-verification.
Verification shall be by analysis, simulation and flight test.
5.3.3.2
Control power for designated secondary flight path controller-verification.
Verification shall be by analysis, simulation and flight test.
5.3.4 Flight path controller characteristics-verification.
Verification shall be by analysis, simulation and test.
5.4
Flying qualities requirements for the longitudinal (speed) axis-verification
5.4.1 Speed response to attitude changes-verification.
Verification shall be by analysis, simulation and flight test.
5.4.1.1
Speed response to attitude changes-relaxation in transonic flight-verification.
Verification shall be by analysis, simulation and flight test.
5.5
Flying qualities requirements for the roll axis-verification
5.5.1
Roll response to roll controller-verification
5.5.1.1
Roll mode-verification.
Verification shall be by analysis, simulation and flight test.
5.5.1.2
Spiral stability-verification.
Verification shall be by analysis, simulation and flight test.
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MIL-STD-1797A
5.5.1.3
Coupled roll-spiral oscillation-verification.
Verification shall be by analysis, simulation and flight test.
5.5.1.4
Roll oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.5.1.5
Roll time delay-verification.
Verification shall be by analysis, simulation and flight test.
5.5.2 Pilot-induced roll oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.5.3 Linearity of roll response to roll controller-verification.
Verification shall be by analysis, simulation and flight test.
5.5.4 Lateral acceleration at the pilot station-verification.
Verification shall be by analysis, simulation and flight test.
5.5.5 Roll characteristics in steady sideslip-verification.
Verification shall be by analysis, simulation and flight test.
5.5.6 Roll axis control for takeoff and landing in crosswinds-verification.
Verification shall be by analysis, simulation and flight test.
5.5.7
Roll axis response to other inputs-verification
5.5.7.1
Roll axis response to augmentation failures-verification.
Verification shall be by analysis, simulation and flight test.
5.5.7.2
Roll axis response to configuration or control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.5.8
Roll axis control power-verification
5.5.8.1
Roll axis response to roll control inputs-verification.
Verification shall be by analysis, simulation and flight test.
5.5.8.2
Roll axis control power in steady sideslips-verification.
Verification shall be by analysis, simulation and flight test.
5.5.8.3
Roll axis control power in crosswinds-verification.
Verification shall be by analysis, simulation and ground and flight test.
5.5.8.4
Roll axis control power for asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.5.8.5
Roll axis control power in dives and pullouts-verification.
Verification shall be by analysis, simulation and flight test.
5.5.8.6
Roll axis control power for asymmetric loading-verification.
Verification shall be by analysis, simulation and flight test.
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MIL-STD-1797A
5.5.9
Roll axis control forces and displacements-verification
5.5.9.1
Roll control displacements-verification.
Verification shall be by inspection or flight test.
5.5.9.2
Roll axis control forces to achieve required roll performance-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.3
Roll axis control sensitivity-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.4
Roll axis control centering and breakout forces-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5
Roll axis control force limits-verification
5.5.9.5.1 Roll axis control force limits in steady turns-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.2 Roll axis control force limits in dives and pullouts-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.3 Roll axis control force limits in crosswinds-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.4 Roll axis control force limits in steady sideslips-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.5 Roll axis control force limits for asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.6 Roll axis control force limits for failures-verification.
Verification shall be by analysis, simulation and flight test.
5.5.9.5.7 Roll axis control force limits for configuration or control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.6
Flying qualities requirements for the yaw axis-verification
5.6.1
Yaw axis response to yaw and side-force controllers-verification
5.6.1.1
Dynamic lateral-directional response-verification.
Verification shall be by analysis, simulation and flight test.
5.6.1.2
Steady sideslips-verification.
Verification shall be by analysis, simulation and flight test.
5.6.1.3
Wings-level turn-verification.
Verification shall be by analysis, simulation and flight test.
5.6.2 Yaw axis response to roll controller-verification.
Verification shall be by analysis, simulation and flight test.
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MIL-STD-1797A
5.6.3 Pilot-induced yaw oscillations-verification.
Verification shall be by analysis, simulation and flight test.
5.6.4 Yaw axis control for takeoff and landing in crosswinds-verification.
Verification shall be by analysis, simulation and flight test.
5.6.5
Yaw axis response to other inputs-verification
5.6.5.1
Yaw axis response to asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.6.5.2
Yaw axis response to failures-verification.
Verification shall be by analysis, simulation and flight test.
5.6.5.3
Yaw axis response to configuration or control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.6.6 Yaw axis control power-verification.
Verification shall be by analysis, simulation and flight test.
5.6.6.1
Yaw axis control power for takeoff, landing, and taxi-verification.
Verification shall be by analysis, simulation and flight test.
5.6.6.2
Yaw axis control power for asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.6.6.3
Yaw axis control power with asymmetric loading-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7 Yaw axis control forces-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.1
Yaw axis control force limits in rolling maneuvers-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.2
Yaw axis control force limits in steady turns-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.3
Yaw axis control force limits during speed changes-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.4
Yaw axis control force limits in crosswinds-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.5
Yaw axis control force limits with asymmetric loading-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.6
Yaw axis control force limits in dives and pullouts-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.7
Yaw axis control force limits for waveoff (go-around)-verification.
Verification shall be by analysis, simulation and flight test.
55
MIL-STD-1797A
5.6.7.8
Yaw axis control force limits for asymmetric thrust during takeoff-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.9
Yaw axis control force limits with flight control failures-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.10 Yaw axis control force limits-control mode change-verification.
Verification shall be by analysis, simulation and flight test.
5.6.7.11 Yaw axis breakout forces-verification.
Verification shall be by test.
5.7
Flying qualities requirements for the lateral flight path axis-verification
5.7.1 Dynamic response for lateral translation-verification.
Verification shall be by analysis, simulation and flight test.
5.8
Flying qualities requirements for combined axes-verification
5.8.1 Cross-axis coupling in roll maneuvers-verification.
Verification shall be by analysis, simulation and flight test.
5.8.2 Crosstalk between pitch and roll controllers-verification.
Verification shall be by analysis, simulation and flight test.
5.8.3 Control harmony-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4
Flight at high angle of attack-verification
5.8.4.1
Warning cues-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.2
Stalls-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.2.1 Stall approach-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.2.2 Stall characteristics-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.2.3 Stall prevention and recovery-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.2.4 One-engine-out stalls-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.3
Post-stall gyrations and spins-verification.
Verification shall be by analysis, simulation and flight test.
5.8.4.3.1 Departure from controlled flight-verification.
Verification shall be by analysis, simulation and __________.
56
MIL-STD-1797A
5.8.4.3.2 Recovery from post-stall gyrations and spins-verification.
Verification shall be by analysis, simulation and __________.
5.9
Flying qualities requirements in atmospheric disturbances-verification
5.9.1 Allowable flying qualities degradations in atmospheric disturbances-verification.
Verification shall be by analysis, simulation and flight test.
5.9.2 Definition of atmospheric disturbance model form-verification.
Verification shall be by analysis and simulation.
5.9.3 Application of disturbance models in analyses-verification.
Verification shall be by analysis and simulation.
6.
NOTES
This section contains information of a general or explanatory nature that may be helpful, but is not
mandatory.
6.1
Intended use.
This standard contains the handling qualities requirements for piloted aircraft and forms one of the bases
for determination by the procuring activity of aircraft acceptability. The standard consists of requirements
in terms of criteria for use in stability and control calculations, analysis of wind tunnel test results,
simulator evaluations, flight testing, etc. The requirements should be met as far as possible by providing
an inherently good basic airframe. Cost, performance, reliability, maintenance, etc. trade-offs are
necessary in determining the proper balance between basic airframe characteristics and augmented
dynamic response characteristics. The contractor should advise the procuring activity of any significant
design penalties which may result from meeting any particular requirement.
6.2
Level definitions.
Part of the intent of 4.1.7.3 and 4.9.1 is to ensure that the probability of encountering significantly
degraded flying qualities because of component or subsystem failures is small. For example, the
probability of encountering very degraded flying qualities (Level 3) must be less than specified values per
flight. To determine the degradation in flying qualities parameters for a given Aircraft Normal State the
following definitions are provided:
a. Level 1 is better than or equal to the Level 1 boundary, or number, specified in section 4.
b. Level 2 is worse than Level 1, but no worse than the Level 2 boundary, or number
c. Level 3 is worse than Level 2, but no worse than the Level 3 boundary, or number.
When a given boundary, or number, is identified as Level 1 and Level 2, this means that flying qualities
outside the boundary conditions shown, or worse than the number given, are at best Level 3 flying
qualities. Also, since Level 1 and Level 2 requirements are the same, flying qualities for Aircraft Normal
States must be within this common boundary, or number, in both the Operational and Service Flight
Envelopes (4.1.6.1). Aircraft Failure States that do not degrade flying qualities beyond this common
boundary are not considered in meeting the requirements of 4.1.7.3. Aircraft Failure States that represent
degradations to Level 3 must, however, be included in the computation of the probability of encountering
Level 3 degradations in both the Operational and Service Flight Envelopes. Again, degradation beyond
the Level 3 boundary is not permitted regardless of component failures.
6.3
Reference documents tree.
The following list of documents comprises the 1st and 2nd tier references. Only 1st tier references are
contractually binding; 2nd tier is for guidance only.
57
MIL-STD-1797A
1st Tier References
2nd Tier References
JAN-1-225
Interference Measurements
JAN-T-781
Terminal; Cable
MIL-F-3541
Fittings, Lubrication
MIL-S-3950
Switches, Toggle
MIL-E-4682
Electron Tubes
MIL-W-5088
Wiring, Aircraft
MIL-E-5272
Environmental Testing
MIL-E-5400
Electronic Equipment
MIL-H-5440
Hydraulic System
MIL-I-6115
Instrument Systems
MIL-1-6181
Interference Control
MIL-L-6880
Lubrication of Aircraft
MIL-E-7080
Electrical Equipment
MIL-M-7969
Motors, Alternating
MIL-A-8064
Actuators and Act Systems
MIL-M-7793
Meter, Time Totaling
MIL-H-8501
Helicopter Flying Qual.
MIL-S-8512
Support Equipment, Aeron
MIL-M-8609
Motors, Direct Current
MIL-D-8706
Data, Design
MIL-F-8785
Flying Qualities
MIL-D-18300
Design Data Requirements
MIL-N-18307
Nomenclature/Nameplates
MIL-E-19600
Electronic Modules
MIL-R-22256
Reliability Requirements
MIL-F-23094
Reliability Assurance
MIL-STD-203
Cockpit Controls
MIL-STD-704
Electnc Power
MS15001
Fittings, Lubrication
MS 15002
Fittings, Lubrication
MIL-C-18244, Control
and Stabilization Systems
MIL-F-87242,
Flight Controls
6.4
Data requirements.
When this standard is used in an acquisition which incorporates a DD Form 1423, Contract Data
Requirement List (CDRL), the data requirements identified below shall be developed as specified by an
approved Data Item Description (DD Form 1664) and delivered in accordance with the approved CDRL
incorporated into the contract. When the provisions of the DoD FAR clause on data requirements
58
MIL-STD-1797A
(currently DoD FAR Supplement 52-227-7031) are invoked and the DD Form 1423 is not used, the data
specified below shall be delivered by the contractor in accordance with the contract or purchase order
requirements. Deliverable data required by this standard is cited in the following paragraphs.
Paragraph No.
Data Requirement Title
Applicable DID No.
Option
(Data item descriptions related to this standard, and identified in section 6 will be approved and listed as
such in DoD 5000.19-L, Vol. II, AMSDL. Copies of data item descriptions required by the contractors in
connection with specific acquisition functions should be obtained from the Naval Publications and Forms
Center or as directed by the contracting officer.)
6.5
Subject term (key word) listing
Aircraft
Airplane design criteria
Design criteria
Handling qualities
Pilot vehicle interface
6.6
Responsible engineering office (REO).
The office responsible for development and technical maintenance of this standard is ASD/ENFTC,
Wright-Patterson AFB, OH 45433-6503; AUTOVON 785-5730, Commercial (513) 255-5730. Any
information relating to Government contracts must be obtained through contracting officers.
6.7
Changes from previous issue.
Marginal notations are not used in this revision to identify changes with respect to the previous issue due
to the extensiveness of the changes.
Custodians:
Army - AV
Navy - AS
Air Force - 11
Preparing activity
Air Force - 11
Project No 15GP-0088
59
MIL-STD-1797A
APPENDIX A
FLYING QUALITIES OF PILOTED AIRCRAFT
HANDBOOK FOR
10.
SCOPE
10.1
Scope.
This appendix provides rationale, guidance, lessons learned, and instructions necessary, to tailor sections
4 and 5 of the basic standard (MIL-STD-1797A) for a specific application.
10.2
Purpose.
This appendix provides information to assist the Government procuring activity in the use of MIL-STD1797A.
10.3
Use.
This appendix is designed to assist the project engineer in tailoring MIL-STD-1797A. The blanks of the
basic standard shall be filled in to meet operational needs of the tailored application
10.4
Format
10.4.1 Requirement/verification identity.
Section 40 of this appendix parallels section 4 and section 5 of the basic standard; paragraph titles and
numbering are in the same sequence. Section 40 provides each requirement (section 4) and associated
verification (section 5) as stated in the basic standard. Both the requirement and verification have
sections for rationale, guidance, and lessons learned.
10.4.2 Requirement/verification package.
Section 40 of this appendix has been arranged so that the requirement and associated verification is a
complete package to permit addition to, or deletion from, the criteria as a single requirement. A
requirement is not specified without an associated verification.
10.5
Responsible engineering office.
The responsible engineering office (REO) for this appendix is ASD/ENFTC, Wright-Patterson AFB OH
45433-6503; AUTOVON 785-5730, Commercial (513)-255-5730.
20.
APPLICABLE DOCUMENTS
20.1
References.
The documents referenced in this appendix are not intended to be applied contractually. Their primary
purpose is to provide background information for the Government engineers responsible for developing
the most appropriate performance values (filling in the blanks) for the requirements contained in the
standard proper.
20.2
Avoidance of tiering.
Should it be determined that the references contained in this appendix are necessary in writing an RFP or
building a contract, excessive tiering shall be avoided by calling out only those portions of the reference
which have direct applicability. It is a goal of the Department of Defense that the practice of referencing
documents in their entirety be eliminated in order to reduce the tiering effect.
20.3
Government documents
SPECIFICATIONS
60
MIL-STD-1797A
APPENDIX A
Military
MIL-C-5011
Charts: Standard Aircraft Characteristics and Performance, Piloted Aircraft
(Fixed Wing)
MIL-D-8708
Demonstration Requirements for Airplanes
MIL-F-8785
Flying Qualities of Piloted Airplanes
MIL-A-8861
Airplane Strength and Rigidity Flight Loads
MIL-F-9490
Flight Control Systems - Design, Installation and Test of Piloted Aircraft,
General Specification for
MIL-F-18372
Flight Control Systems:
Specification for)
MIL-S-25015
Spinning Requirements for Airplanes
MIL-W-25140
Weight and Balance Control System (for Airplanes and Rotorcraft)
MIL-F-83300
Flying Qualities of Piloted V/STOL Aircraft
MIL-S-83691
Stall/post-stall/spin Flight Test Demonstration Requirements for Airplanes
AFGS-87221
Aircraft Structures, General Specification for
Navy BuAer SR119B/USAF
C-1815B
Flying Qualities of Piloted Airplanes
Design, Installation and Test of, Aircraft (General
STANDARDS
Military
MIL-STD-756
Reliability Modeling and Prediction
MIL-STD-785
Reliability Program for Systems and Equipment Development and Production
MIL-STD-882
System Safety Program Requirements
MIL-STD-1629
Procedures for Performing a Failure Mode, Effects and Criticality Analysis
HANDBOOKS
Military
MIL-HDBK-217
Reliability Prediction of Electronic Equipment
MIL-HDBK-244
Guide to Aircraft/Stores Compatibility
REPORTS
Navy Rpt No.
SA-C7R-75
First Interim Report, Flying Qualities Technical Evaluation of the F-14A
Airplane; Humphrey, M. J.; November 1975 (declassified 31 December 1981)
Navy Rpt No.
SA-14R-81
Navy Evaluation of the F/A-18A Airplane with Roll Rate Improvements
Incorporated; Copeland, W., K. Grubbs, et al; March 1981
ASD-TDR-61-362
Fixed-Base and In-Flight Simulation of Longitudinal and Lateral-Directional
Handling Qualities for Piloted Re-entry Vehicles; Kidd, E. A. and R. P. Harper;
February 1964
61
MIL-STD-1797A
APPENDIX A
ASD-TDR-62-507
Handling Qualities in Single-Loop Roll Tracking Tasks: Theory and Simulator
Experiments; Durand, T. S. and H. R. Jex; November 1962
ASD-TDR-63-399
Fixed-Base Simulator Investigation of the Effects of Lα, and True Speed on
Pilot Opinion of Longitudinal Flying Qualities; Chalk, C. R.; November 1963
ASD-TR-72-48
Criteria for Predicting Spin Susceptibility of Fighter-Type Aircraft; Weissman,
R.; June 1972
ASD-TR-78-13
USAF Flying Qualities Requirements for a STOL Transport; Gerken, G.; May
1979
WADC-TR-52-298
Artificial Stability Flight Tests of the XF-88A Airplane; Moore, N. B.; July 1954
WADC-TR-54-594
Flight Evaluations of Variable Short Period and Phugoid Characteristics in a B26; Newell, F. d. and G. Campbell; December 1954
WADC-TR-55-299
Flight Evaluations of Various Longitudinal Handling Qualities in a VariableStability Jet Fighter; Harper, R. P., Jr.; July 1955
WADC-TR-56-258
Flight Evaluations in Variable-Stability Airplanes of Elevator Control Motion
Gradients for High-Speed Bombers; Harper, R. P., Jr ; November 1956
WADC-TR-57-520
Human Pilot Dynamic Response; Seckel, E., 1. A. M. Hall, et al.; August 1958
WADC-TR-57-719
Part II
Additional Flight Evaluations for Various Longitudinal Handling Qualities in a
Variable-Stability Jet Fighter, Chalk, C. R.; July 1958
WADC-TR-58-82
Approximate Airframe Transfer Functions and Application to Single Sensor
Control Systems; Ashkenas, I. L. and D. T. McRuer; June 1958
WADC-TR-59-135
The Determination of Lateral Handling Quality Requirements from AirframeHuman Pilot System Studies; Ashkenas, I. L. and D. T. McRuer; June 1959
WADD-TR-61-147
In-Flight Simulation of the Lateral-Directional Handling Qualities of Entry
Vehicles; Harper, R. P., Jr.; November 1961
AFFDL-TR-65-15
Human Pilot Dynamics in Compensatory Systems-Theory, Models, and
Experiments with Controlled Element and Forcing Function Variations,
McRuer, D., D. Graham, et al.; July 1965
AFFDL-TR-65-39
Ground Simulator Evaluations of Coupled Roll-Spiral Mode Effects on Aircraft
Handling Qualities; Newell, F. D.; March 1965
AFFDL-TR-65-138
A Study of Conventional Airplane Handling Qualities Requirements. Part 1:
Roll Handling Qualities; Ashkenas, I. L.; November 1965
AFFDL-TR-65-198
A Handling Qualities Theory for Precise Flight-Path Control; Bihrie, Jr.; June
1966
62
MIL-STD-1797A
APPENDIX A
AFFDL-TR-65-210
Simulated Landing Approaches of an Unaugmented C-5A Configuration;
Newell, F. D.. M. L. E. Parrag and G. Bull; December 1965
AFFDL-TR-65-218
Estimation of Flying Qualities of Piloted Airplanes; Woodcock, R. J. and D. E.
Drake; April 1966
AFFDL-TR-65-227
Supersonic Transport Handling Characteristics During Approach and Landing
Flight Regimes; Klein, R. H., R. B. Archer and D. W. Lew; December 1965
AFFDL-TR-66-2
Flight Evaluation of Various Phugoid and I/Th, Values for the Landing
Approach Task; Chalk, C. R.; February 1966
AFFDL-TR-66-148
Flying Qualities Conference, Wright-Patterson Air Force Base, Ohio, 5 and 6
April, 1966; December 1966
AFFDL-TR-66-163
Flight Investigation of Longitudinal Short Period Frequency Requirements and
PIO Tendencies; DiFranco, D. A.; June 1967
AFFDL-TR-67-2
Analysis of Several Handling Quality Topics Pertinent to Advanced Manned
Aircraft; Stapleford, R. L. and J. A. Tennant; June 1967
AFFDL-TR-67-19
Pilot Evaluations in a Ground Simulator of the Effects of Elevator Control
System Dynamics in Fighter Aircraft; Keith, L. A., R. R. Richard and G. J.
Marrett; December 1968
AFFDL-TR-67-51
In-Flight Simulation and Pilot Evaluation of Selected Landing Approach
Handling Qualities of a Large Logistics Transport Airplane; Rhoads, D W.; July
1967
AFFDL-TR-67-98
In-Flight Evaluation of Lateral-Directional Handling Qualities for the Fighter
Mission; Mecker, J. I. and G. W. Hall; October 1967
AFFDL-TR-68-85
Investigation of the Effects of Gusts on V/STOL Craft in Transit)on and Hover;
Skelton, G. B.; October 1968
AFFDL-TR-68-90
In-Flight Investigation of the Effects of Higher Order Control System Dynamics
on Longitudinal Flying Qualities; DiFranco, D.; August 1968
AFFDL-TR-68-91
In-Flight Investigation of Longitudinal Short-Period Handling Characteristics of
Wheel-Controlled Airplanes; Hall, G. W.; August 196,q
AFFDL-TR-69-3
In-Flight Investigation of the Effect on PIO of Control System Nonlinearities,
Pitch Acceleration and Normal Acceleration Bobweights-, Newell, F. D. and R.
Wasserman; May 1969
AFFDL-TR-69-41
A Flight Investigation of Lateral-Directional Handling Qualities for V/STOL
Aircraft in Low Speed Maneuvering Flight; Doetsch, K. H. Jr., D. G. Gould and
D. M. McGregor; March 1976
AFFDL-TR-69-67
A Non-Gaussian Turbulence Simulation; Reeves, P. M.; November 1969
AFFDL-TR-69-72
Background Information and User Guide for MIL-F-8785B(ASG), "Military
Specification - Flying Qualities of Piloted Airplanes"; Chalk, C R., T. P. Neal, et
al.; August 1969
63
MIL-STD-1797A
APPENDIX A
AFFDL-TR-70-74
Vols I and II
An In-Flight Investigation to Develop Control System Design Criteria for Fighter
Airplanes; Neal, T. P. and Rogers E. Smith, December 1970
AFFDL-TR-70-145
An In-Flight Investigation of Lateral-Directional Dynamics for the Landing
Approach; Hall, G. W. and E. M. Boothe; October 1971
AFFDL-TR-70-155
Validation of the Flying Qualities Requirements of MIL-F-00878SA(USAF);
Brady, C. C. and J. Hodgkinson; January 1971
AFFDL-TR-71-134
Validation of the Flying Qualities Requirements of MIL-F-8785B(ASG),
Kandalaft, R. N.; September 1971
AFFDL-TR-71-164
Vol I
In-Flight Investigation of an Unaugmented Class III Airplane in the Landing
Approach Task. Phase I: Lateral-Directional Study; Wasserman, R., F. F.
Eckhart and H. J. Ledder; January 1972
AFFDL-TR-72-36
Evaluation of Lateral-Directional Handling Qualities and Roll-Sideslip Coupling
of Fighter Class Airplanes; Boothe, E. M. and M. L. Parrag, May 1972
AFFDL-TR-72-41
Revisions to MIL-F-8785B(ASG) Proposed by Cornell Aeronautical Laboratory
Under Contract F33615-71-C-1254; Chalk, C. R., D A. DiFranco, et al.; April
1973
AFFDL-TR-72-141
Vol I
Validation of the Flying Qualities Requirements of MIL-F-8785B(ASG) Using
the P-3B Airplane; Richards, R. B., D. L. Green and J. C. Rennie, November
1973
AFFDL-TR-72-143
In-Flight Simulation of Minimum Longitudinal Stability for Large Delta-Wing
Transports in Landing Approach and Touchdown. Vol 1: Technical Results;
Wasserman, R. and J. F. Mitchell; February 1973
AFFDL-TR-73-76
Recommended Revisions to Selected Portions of MIL-F-8785B(ASG) and
Background Data; Ashkenas, I. L., R. H. Hoh and S. J. Craig; August 1973
AFFDL-TR-74-9
A Two-Phase Investigation of Longitudinal Flying Qualities for Fighters;
Boothe, E. M., R. T. N. Chen and C. R. Chalk; April 1974
AFFDL-TR-74-61
Investigation of Flying Qualities of Military Aircraft at High Angles of Attack. Vol
1: Technical Results; Johnston, D. E., I. L. Ashkenas and J. R. Hogge; June
1974
AFFDL-TR-74-130
(2 Vols)
Extension of the Method for Predicting Six-Degree-of-Freedom Store
Separation Trajectories at Speeds Up to the Critical Speed to Include A
Fuselage with Noncircular Cross Section; Dillenius, M. F. E., F. K. Goodwin
and J. N. Nielsen; November 1974
AFFDL-TR-75-3
Evaluation of the Flying Qualities Requirements of MIL-F-8785B(ASG) Using
the C-5A Airplane-, Silvers, C. L. and C. C. Withers; March 1975
AFFDL-TR-76-78
Direct Side Force Control Criteria for Dive Bombing. Vol 1: Summary Vol II:
Analysis and Results; Brulle, R. V., W. A. Moran and R. G Marsh; September
1976
64
MIL-STD-1797A
APPENDIX A
AFFDL-TR-77-57
A Theory for Longitudinal Short-Period Pilot Induced Oscillations; Smith, Ralph
H.; June 1977
AFFDL-TR-78-9
Fighter CCV Phase IV Report, Vol II: Flight Test Data Evaluation. Vol III: Test
Phase Data Summary, Parts I and 2; McAllister, J. D., et al.; February 1978
AFFDL-TR-78-122
Effects of Control System Dynamics on Fighter Approach and Landing
Longitudinal Flying Qualities (Volume I); Smith, Rogers E.; March 1978
AFFDL-TR-78-171
Proceedings of AFFDL Flying Qualities Symposium Held at Wright State
University 12-15 September, 1978; Black, G. T., Moorhouse, D. J., et al.,
compilers; December 1978:
“Task-Oriented Flying Qualities for Air-to-Ground Gun Attack;" Brandeau, G.
“B-1 Experience Related to MIL-F-8785B and Proposed Revisions;" Campbell,
J. E.
"An Approach to Simplify the Specification of Low-Speed Maneuvering Pitch
Control Force;" Cichy, D. R.
"High Angle of Attack Flying Qualities and Departure Criteria Development;"
Hellman, G. K. and R. B. Crombie
"Northrop Review of MIL-F-8785B Proposed Revision;" Lockenour, J
“Evaluation of Selected Class III Requirements of MIL-F-8785B(ASG),
“Flying Qualities of Piloted Airplanes;'" Withers, C. C.
"Discussion and Status of the Proposed Revision (1978) to MIL-F-8785B;"
Moorhouse, D. J., R. J. Woodcock and T. P. Sweeney
AFFDL-TR-79-3126
Flight Qualities Design Requirements for Sidestick Controllers; Black, G. T.
and D. J. Moorhouse; October 1979
AFWAL-TR-80-3032
Prediction of Supersonic Store Separation Characteristics Including Fuselage
and Stores of Noncircular Cross Section (4 volumes); Goodwin, F. K., M. F. E.
Dillenius and J. Mullen. Jr.; November 1980
AFWAL-TR-80-3060
Simulation Analysis: Unorthodox Control Force Fighter Aircraft, Vol II: Detailed
Summary; Mitchell, A. L., et al.; April 1980
AFWAL-TR-80-3067
Flying Qualities Design Criteria: Proceedings of AFFDL Flying Qualities
Symposium Held at Wright-Patterson Air Force Base in October 1979.
Crombie, R. B. and D. J. Moorhouse, compilers; May 1980
AFWAL-TR-80-3141
Investigation of High-Angle-of-Attack Maneuvering - Limiting Factors, Part 1:
Analysis and Simulation; Johnston, D. E., D. G. Mitchell and T T. Myers;
December 1980
AFWAL-TR-81-3027
Development of Handling Quality Criteria for Aircraft with Independent Control
of Six-Degrees-of-Freedom; Hoh, R. H., T. T. Myers, et al.; April 1981
65
MIL-STD-1797A
APPENDIX A
AFWAL-TR-81-3108
Investigation of High AOA Flying Qualities and Design Guides; Johnston, D. E.
and R. K. Heffley; December 1981
AFWAL-TR-81-3109
Background Information and User Guide for MIL-F-8785C. Military
Specification - Flying Qualities of Piloted Airplanes; Moorhouse, D J. and R. J.
Woodcock; September 1981
AFWAL-TR-81-3116
Equivalent System Verification and Evaluation of Augmentation Effects on
Fighter Approach and Landing Flying Qualities; Smith, Rogers E.; September
1981
AFWAL-TR-81-3118
In-Flight Investigation of Large Airplane Flying Qualities for Approach and
Landing; Weingarten, N. C. and C. R. Chalk; September 1981
AFWAL-TR-81-3171
Lateral Flying Qualities of Highly Augmented Fighter Aircraft, Vols. I and 11;
Monegan, S. J., Rogers E. Smith and R. E. Bailey; June 1982
AFWAL-TR-82-3014
Proposed Revisions to MIL-F-8785C Related to Flight Safety of Augmented
Aircraft, 3 Vols.; Schuler, J. M. and M A. Dahl, April 1982
AFWAL-TR-82-3064
Design Criteria for the Future of Flight Controls, Proceedings of the Flight
Dynamics Laboratory Flying Qualities and Flight Control Symposium, 2-5
March, 1982; Fuller, S. G. and Potts, D. W., compilers; July 1982
AFWAL-TR-82-3081
Proposed MIL Standard and Handbook - Flying Qualities of Air Vehicles, Vol II:
Proposed MIL Handbook; Hoh, R. H., Mitchell, D. G., et al November 1982
AFWAL-TR-83-3015
Suggested Revisions to MIL-F-8785C for Large Class III Aircraft; Nleyer, R. T.,
et al.; February 1983
AFFDL-FGCTM-71-7
Validation of the Handing Qualities Degradation Probabilities of MIL-F008785A Using F-4C Air Force Manual 66-1 Maintenance Data, Ullman, Lt., T.
Calanducci, and Lt. Linck; August 1971
AFAMRL-TR-73-78
Manual Control Performance and Dynamic Response During Sinusoidal
Vibration; Allen, R. Wade, Henry R. Jex, and Raymond E. Magdaleno, October
1973
AFAMRL-TR-81-39
Male and Female Strength Capabilities for Operating Aircraft Controls;
McDaniel, Joe W.; March 1981
AFFTC-SD-69-5
A-7D Stability and Control Military Preliminary Evaluations (Phase IA and IB);
Gobert, Don 0. and William T. Twinting; April 1969
AFFTC-TD-75-1
Tracking Test Techniques for Handling Qualities Evaluation; Twisdale, T R.
and D. L. Franklin; May 1975
AFFTC-TR-75-15
Flying Qualities Evaluation of the YF-16 Prototype Lightweight Fighter, Eggers,
James A. and William F. Bryant, Jr.; July 1975
AFFTC-TR-75-32
F-15A Approach-to-Stall/Stall/Post-Stall Evaluation; Wilson, Donald B and
Charles A. Winters; January 1976
AFFTC-TR-76-15
Flight Test Development and Evaluation of a Multimode Digital Flight Control
System Implemented in an A-7D (DIGITAC), Damman,
66
MIL-STD-1797A
APPENDIX A
Lawrence, Robert Kennington, Paul Kirsten, Ronald Grabe, and Patrick Long;
June 1976
AFFTC-TR-77-27
System Identification from Tracking (SIFT), a New Technique for Handling
Qualities Test and Evaluation (Initial Report); Twisdale, T. R and T. A. Ashurst;
November 1977
AFFTC-TR-79-2
Flying Qualities and Flight Control System Evaluation of the B-1 Strategic
Bomber; Ross, Jerry L., Page G. McGirr, and Otto J. Waniczek, Jr.; May 1979
AFFTC-TR-79-10
F-16A/B Flying Qualities Full-Scale Development Test and Evaluation; Pape,
James A. and Michael P. Garland; September 1979
AFFTC-TR-79-18
F-16A/B High Angle of Attack Evaluation; Wilson, Donald B. and Robert C.
Ettinger; October 1979
AFFTC-TR-80-23
F-15C Flying Qualities Air Force Development Test and Evaluation; Shaner,
Keith L. and Robert W. Barham; November 1980
AFFTC-TR-80-29
F-16 Flying Qualities with External Stores; Garland, Michael P., Michael K.
Nelson, and Richard C. Patterson; February 1981
FDL-TDR-64-60
Flight Evaluation of Various Short Period Dynamics at Four Drag
Configurations for the Landing Approach Task; Chalk, C. R., October 1964;
Chalk, Charles R.; October 1964
FTC-TR-66-24
Frequency Response Method of Determining Aircraft Longitudinal Short Period
Stability and Control System Characteristics in Flight; Klung, H A., Jr.; August
1966
FTC-TR-67-19
Evaluation of Longitudinal Control Feel System Modifications Proposed for
USAF F/RF-4 Aircraft ; Keith, L. A., R. R. Richard, and G J. Marrett, December
1968
FTC-TD-72-1
Development and Evaluation of the TWeaD II Flight Control Augmentation
System; Carleton, David L., Richard E. Lawyer, and Cecil W. Powell;
November 1972
FrC-TD-73-2
Background Information and User Guide for MIL-S-83691; Sharp, Patrick S.
and Collet E. McElroy, March 1974
FTC-TR-73-32
Air Force Evaluation of the Fly-by-Wire Portion of the Survivable Flight Control
System Advanced Development Program; Majoros, Robert L.; August 1973
FTC-TIH-79-2
USAF Test Pilot School, Flying Qualities Handbook, Flying Qualities Theory
and Flight Test Techniques; November 1979
USNTPS-FTM-103
Fixed Wing Stability and Control, Theory and Flight Techniques; 1 November
1981
FAA FAR Part 23
Airworthiness Standards: Normal, Utility, and Acrobatic Category Airplanes;
June 1974
67
MIL-STD-1797A
APPENDIX A
FAA FAR Part 25
Airworthiness Standards: Transport Category Airplanes; June 1974
FAA-ADS-69-13
An In-Flight Investigation of Lateral-Directional Dynamics for Cruising Flight;
Hall, G. W.; December 1969
FAA-RD-70-61
A Flight Simulator Study of STOL Transport Lateral Control Characteristics;
Drake, Douglas E., Robert A. Berg, Gary L. Teper, and W. Allen Shirley;
September 1970
FAA-RD-70-65
Flying Qualities of Small General Aviation Airplanes. Part 2: The Influence of
Roll Control Sensitivity Roll Damping, Dutch-Roll Excitation, and Spiral
Stability; Ellis, David R; April 1970
FAA-RD-74-206
Wind Models for Flight Simulator Certification of Landing and Approach
Guidance and Control Systems; Barr, Neal M., Dagfinn Gangsaas, and Dwight
R. Schaeffer; December 1974
FAA-RD-75-123
Identification of Minimum Acceptable Characteristics for Manual STOL Flight
Path Control; Hoh, Roger H., Samuel J. Craig, and Irving L. Ashkenas; June
1976
FAA-RD-77-25
A Study of Lightplane Stall Avoidance and Suppression; Ellis, David R.;
February 1977
FAA-RD-77-36
Wind Shear Modeling for Aircraft Hazard Definition; Frost, Walter and Dennis
W. Camp; March 1977
FAA-RD-77-173
Proceedings of the First Annual Meteorological and Environmental Inputs to
Aviation Systems Workshop. "Wind Models for Flight Simulator Certification of
Landing and Approach Guidance and Control Systems", Schaeffer, Dwight R.;
March 1977
FAA-RD-78-7
Simulation and Analysis of Wind Shear Hazard; Lehman, John M., Robert K.
Heffley, and Warren F. Clement; December 1977
FAA-RD-79-59
Powered-Lift Aircraft Handling Qualities in the Presence of Naturally-Occurring
and Computer-Generated Atmospheric Disturbances; Jewell, Wayne F.,
Warren F. Clement, Thomas C. West, and S. R. M. Sinclair; May 1979
FAA-RD-79-84
Piloted Flight Simulation Study of Low-Level Wind Shear, Phase 4; Foy, W. H.
and W. B. Gartner; March 1979
FAA Advisory
Circular
AC25.253-1A
High-Speed Characteristics; 24 November 1965
DOT/FAA/CT-82/
130-II
Flying Qualities of Relaxed Static Stability Aircraft, Vol II; McRuer, D. T and T.
T. Myers; September 1982
NACA Memo
Rpt L6E20
Flight Investigation to Improve the Dynamic Longitudinal Stability and ControlFeel Characteristics of the P-63A-1 Airplane with Closely Balanced
Experimental Elevators, Johnson, Harold I.; July 1946
NASA Memo
1-29-59A
A Pilot Opinion Study of Lateral Control Requirements for Fighter- Type
Aircraft; Creer, Brent Y., John D. Stewart, Robert B. Merrick, and Fred J.
Drinkwater III; March 1959
68
MIL-STD-1797A
APPENDIX A
NASA Memo
12-10-58A
A Flight Investigation to Determine the Lateral Oscillatory Damping Acceptable
for an Airplane in the Landing Approach; McNeill, Walter E. and Richard F.
Vomaske; February 1959
NASA-CP-2028
Proceedings of the First Annual Meteorological and Environmental Inputs to
Aviation Systems Workshop, 'Wind Models for Flight Simulator Certification of
Landing and Approach Guidance and Control Systems"; Schaeffer, Dwight R.;
March 1977
NASA-CR-239
Development of Satisfactory Lateral-Directional Handling Qualities in the
Landing Approach; Stapleford, Robert L., Donald E. Johnston, Gary L. Teper,
and David H. Weir; July 1965
NASA-CR-635
In-Flight and Ground Based Simulation of Handling Qualities of Very Large
Airplanes in Landing Approach; Condit, Philip M., Laddie G. Kimbrel, and
Robert G. Root; October 1966
NASA-CR-778
Evaluation of Lateral-Directional Handling Qualities of Piloted Re-EntryVehicles Utilizing Fixed-Base and In-Flight Evaluations; Meeker, J I May 1967
NASA-CR-2017
Handling Qualities Criteria for the Space Shuttle Orbiter During the Terminal
Phase of Flight; Stapleford, Robert L., Richard H. Klein, and Roger H. Hoh;
April 1972
NASA-CR-2451
Non-Gaussian Atmospheric Turbulence Model for Use in Flight Simulators;
Reeves, P. M., G. S. Campbell, V. M. Ganzer, and R. G. Joppa; September
1974
NASA-CR-2677
Manual and Automatic Flight Control During Severe Turbulence Penetration;
Johnston, Donald E., Richard H. Klein, and Roger H. Hoh, April 1976
NASA-CR-152064
Investigation of the Vulnerability of Powered Lift STOLs to Wind Shear; Hoh,
Roger H. and Wayne F. Jewell; October 1976
NASA-CR-152139
Study of a Safety Margin System for Powered-Lift STOL Aircraft; Heffley,
Robert K. and Wayne F. Jewell; May 1978
NASA-CR-152194
A Study of Key Features of the RAE Atmospheric Turbulence Model; Jewell,
Wayne F. and Robert K. Heffley; October 1978
NASA-CR-159059
An Investigation of Low-Speed Lateral Acceleration Characteristics of
Supersonic Cruise Transports Using the Total In-Flight Simulator (TIFS),
Weingarten, N. C.; July 1979
NASA-CR-159236
Calspan Recommendations for SCR Flying Qualities Design Criteria-, Chalk,
C. R.; April 1980
NASA-CR-163108
Analyses of Shuttle Orbiter Approach and Landing Conditions; Teper, Gary L.,
Richard J. DiMarco, Irving L. Ashkenas, and Roger H. Hoh; July 1981
NASA-CR-172491
Pitch Rate Flight Control Systems in the Flared Landing Task and Design
Criteria Development; Berthe, C. J., C. R. Chalk, and S. Sarrafian
69
MIL-STD-1797A
APPENDIX A
NASA-CR-177331
Mission-Oriented Requirements for Updating MIL-H-8501, Vols. I and II;
Clement, W. F., et al.; January 1985
NASA-TM-86728
Application of Frequency Domain Handling Qualities Criteria to the
Longitudinal Landing Task; Sarrafian, S. K. and B. G. Powers; August 1985
NASA-TM-X-62
Motion Simulator Study of Longitudinal Stability Requirements for Large Delta
Wing Transport Airplanes During Approach and Landing with Stability
Augmentation Systems Failed; Snyder, C. T., E. B. Fry, et al., December 1972
NASA-TM-X-1584
A Review of Transport Handling-Qualities Criteria in Terms of Preliminary, XB70 Flight Experience; Powers, Bruce G.; May 1968
NASA-TN-D-173
Flight Investigation of Automatic Stabilization of an Airplane Having Static
Longitudinal Instability; Russell, Walter R., S. A. Sjoberg, and William L Alford;
December 1959
NASA-TN-D-211
Flight Investigation of Pilot's Ability to Control an Airplane Having Positive and
Negative Static Longitudinal Stability Coupled with Various Effective Lift-Curve
Slopes; Brissenden, Roy F., William L Alford, and Donald L. Mallick; February
1960
NASA-TN-D-746
Flight Controllability Limits and Related Human Transfer Functions as
Determined from Simulator and Flight Tests, Taylor, Lawrence W. and Richard
E. Day; May 1961
NASA-TN-D-779
Flight Investigation Using Variable-Stability Airplanes of Minimum Stability,
Requirements for High-Speed, High-Altitude Vehicles; McFadden, Norman M.,
Richard F. Vomaske, and Donovan R Heinle; April 1961
NASA-TN-D-792
Attitude Control Requirements for Hovering Control Through the Use of a
Piloted Flight Simulator; Faye, A. E., Jr.; April 1961
NASA-TN-D-1141
The Effect of Lateral-Directional Control Coupling on Pilot Control of an
Airplane as Determined in Flight and in a Fixed-Base Flight Simulator;
Vomaske, Richard F., Melvin Sadoff, and Fred J. Drinkwater III; November
1961
NASA-TN-D-1328
A Flight Determination of the Attitude Control Power and Damping
Requirements for a Visual Hovering Task in the Variable Stability and Control
X-14A Research Vehicle; Rolls, L. S. and F. J. Drinkwater; May 1962
NASA-TN-D-1552
A Study of a Pilot's Ability to Control During Simulated Augmentation System
Failures; Sadoff, Melvin; November 1962
NASA-TN-D-1888
A Preliminary Study of Handling-Qualities Requirements of Supersonic
Transports in High-Speed Cruising Flight Using Piloted Simulators; White,
Maurice D., Richard F. Vomaske, Walter E. McNeill, and George E Cooper;
May 1963
NASA-TN-D-2251
A Piloted Simulator Study of Longitudinal Handling Qualities of Supersonic
Transport in the Landing Maneuver; Bray, Richard S., April 1964
70
MIL-STD-1797A
APPENDIX A
NASA-TN-D-3726
An Evaluation of the Handling Qualities of Seven General-Aviation Aircraft;
Barber, Marvin R., Charles K. Jones, Thomas R. Sisk, and Fred W. Haise;
November 1966
NASA-TN-D-3910
A Simulator and Flight Study of Yaw Coupling in Turning Maneuvers of Large
Transport Aircraft; McNeill, W. E. and R. C. Innis; May 1967
NASA-TN-D-3971
Determination of Flight Characteristics of Supersonic Transports During the
Landing Approach with a Large Jet Transport In-Flight Simulator; June 1967
NASA-TN-D-5153
The Use of Pilot Rating in the Evaluation of Aircraft Handling Qualities;
Cooper, G. E. and Harper, R. P., Jr.; April 1969
NASA-TN-D-5466
Simulator Study of Coupled Roll-Spiral Mode Effects on Lateral-Directional
Handling Qualities; Grantham, W. D., F. L. Moore, P. L. Deal, and J. M. Patton,
Jr.; March 1970
NASA-TN-D-5957
Flight Investigation of the Roll Requirements for Transport Airplanes in
Cruising Flight; Holleman, Euclid C.; September 1970
NASA-TN-D-6496
Analysis of a Coupled Roll-Spiral-Mode, Pilot-Induced Oscillation Experienced
with the M2-F2 Lifting Body; Kempel, R. W.; September 1971
NASA-TN-D-6811
In-Flight Pilot Evaluations of the Flying Qualities of a Four-Engine Jet
Transport; Holleman, Euclid C. and Glenn B. Gilyard; May 1972
NASA-TN-D-7703
Flight Investigation of Advanced Control Systems and Displays for a General
Aviation Airplane; Loschke, Paul C., Marvin R. Barber, Einar K Enevoldson,
and Thomas C. McMurtry; June 1974
NASA-TP-1368
Flight Comparison of the Transonic Agility of the F-111A Airplane and the F111 Supercritical Wing Airplane; Friend, Edward L. and Glenn M. Sakamoto;
December 1978
NASA-TR-R-199
Dynamic Response of Airplanes to Atmospheric Turbulence Including Flight
Data on Input and Response; Houbolt, John C., Roy Steiner, and Kermit G.
Pratt; June 1964
NADC-ED-6282
Proposal for a Revised Military Specification, 'Flying Qualities of Piloted
Airplanes' (MIL-F-8785ASG) with Substantiating Text; Mazza, C. J., William
Becker, et al.; 22 July 1963
NADC-76154-30
"Design Charts and Boundaries for Identifying Departure Resistant Fighter
Configurations;" Bihrie, W., Jr. and Barnhart, B.; July 1978
NADC-77052-30
Development of VTOL Flying Qualities Criteria for Low Speed and Hover, Hoh,
Roger H. and Irving L. Ashkenas; December 1979
NADC-78182-60
Development and Analysis of a CVA and a 1052 Class Fast Frigate Air Wake
Model; Nave, Ronald L.; September 1978
NADC-81186-60
The Control Anticipation Parameter for Augmented Aircraft, Bischoff, D E.;
May 1981
71
MIL-STD-1797A
APPENDIX A
NADC-85091-60
“Investigation of Departure Susceptibility Criteria Using the Dynamic Flight
Simulator;" Rhodeside, G.; June 1985
20.4 Nongovernment documents
AIAA Paper 64-353
Jet Transport Operation in Turbulence; Soderlind, Paul A.; July 1964
AIAA Paper 69-898
Summary and Interpretation of Recent Longitudinal Flying Qualities
Results; Ashkenas, I. L.; August 1969
AIAA Paper 75-985
In-Flight Simulation-of the Light Weight Fighters; Hall, G. W. and R. P.
Harper; August 1975
AIAA Paper 77-1119
Direct-Force Flight-Path Control--the New Way to Fly; Watson, John H. and
Jack D. McAllister; August 1977
AIAA Paper 77-1122
Equivalent System Approaches to Handling Qualities Analysis and Design
Problems in Augmented Aircraft; Hodgkinson, J. and W. J. LaManna, 8-10
August 1977
AIAA Paper 77-1145
A Study of Key Features of Random Atmospheric Disturbance Models for
the Approach Flight Phase; Heffley, Robert K.; August 1977
AIAA Paper 78-1500
Rolling Tail Design and Behavior as Affected by Actuator Hinge Moment,
Ball, J. M.; August 1978
AIAA Paper 19-1781
Initial Results of an Inflight, Simulation of Augmented Dynamics in Fighter
Approach and Landing; Hodgkinson, J. and K. A. Johnston; 6-8 August
1979
AIAA Paper 79-1962
Flight Tests of a Microprocessor Control System; Stengel, R F. and G. E.
Miller; October 1979
AIAA Paper 80-0703
Review of Nonstationary Gust-Responses of Flight Vehicles, Gaonkar, G.
H.; July 1980
AIAA Paper 80-1611-CP
Flight Evaluation of Augmented Fighter Aircraft; Hodgkinson, J. and R. C.
Snyder; 11-13 August 1980
AIAA Paper 80-1626-CP
A Summary of an In-Flight Evaluation of Control System Pure Time Delays
During Landing Using the F-8 DFBW Airplane; Berry, D. T., B G. Powers,
K. J. Szalai, and R. J. Wilson; 11-13 August 1980
AIAA Paper 80-1627-CP
Low Order Equivalent Models of Highly Augmented Aircraft Determined
from Flight Data Using Maximum Likelihood Estimation; Shafer, M. F; 11-13
August 1980
AIAA Paper 80-1628-CP
Handling Qualities Criteria for Wing-Level-Turn Maneuvering During an Air
to Ground Delivery; Sammonds, R. I. and J. W. Bunnell, Jr.-, August 1980
AIAA Paper 80-1633
Identification of Flexible Aircraft from Flight Data; Eulrick, B. J. and E. D.
Rynaski; August 1980
AIAA Paper 80-1836
The Turbulent Wind and Its Effect on Flight; Etkin, B.; August 1980
72
MIL-STD-1797A
APPENDIX A
AIAA Paper 81-0302
Atmospheric Disturbance Models and Requirements for the Flying Qualities
Military Standard and Handbook; Heffley, R. K., W. F. Jewell, R. H. Hoh,
and D. J. Moorhouse; January 1981
AIAA Paper 87-2561
Analysis and Application of Aircraft Departure Prediction Criteria to the AV8B Harrier II; Tinger, H.L.; August 1987
SAE ARP 842B
Design Objectives for Flying Qualities of Civil Transport Aircraft
Delft Univ of Tech
Memo M-304
Non-Gaussian Structure of the Simulated Turbulent Environment in Piloted
Flight Simulation; van de Moeskijk, G. A. J.; April 1978
Princeton Univ Rpt 604
A Study of Pilot-Induced Lateral-Directional Instabilities; Caporali, R. L., J.
P. Lamers, and J. R. Totten; May 1962
Princeton Univ Rpt 727
Lateral-Directional Flying Qualities for Power Approach; Seckel, E., G. E.
Miller, and W. B. Nixon; September 1966
Princeton Univ Rpt 777
Comparative Flight Evaluation of Longitudinal Handling Qualities in Carrier
Approach; Eney, J. A.; May 1966
Princeton Univ Rpt 797
Lateral-Directional Flying Qualities for Power Approach: Influence of Dutch
Roll Frequency; Seckel, E., J. A. Franklin, and G E. Miller, September 1967
Stanford Univ
SUDAAR No. 489
Wind Modeling and Lateral Aircraft Control for Automatic Landing; Holley,
William E. and Arthur E. Bryson; January 1975
ARC R&M No. 917
Preliminary Report on Behavior of Aeroplanes When Flying Inverted with
Special Reference to Some Accidents on "A"; O'Gorman, Mervyn,
Chairman, Accidents Committee; January 1919
ESDU Item No. 74031
Characteristics of Atmospheric Turbulence Near the Ground. Part II: Single
Point Data for Strong Winds (Neutral Atmosphere), October 1974
ESDU Item No. 75001
Characteristics of Atmospheric Turbulence Near the Ground. Part III
Variations in Space and Time for Strong Winds (Neutral Atmosphere), July
1975
IAS Paper 60-18
Development of Lateral-Directional Flying Qualities Criteria for Supersonic
Vehicles, Based on a Stationary Flight Simulator Study; Crone, R M. and R.
C. A'Harrah; January 1960
ICAS-86-5.3.4
Handling Qualities for Unstable Combat Aircraft; Gibson, J. C.; September
1986
MDC Rpt A5596
Flying Qualities Analysis of an In-Flight Simulation of High Order Control
System Effects on Fighter Aircraft Approach and Landing-, Johnston, K A.
and J. Hodgkinson, 22 December 1978
MDC Rpt A6792
Definition of Acceptable Levels of Mismatch for Equivalent Systems of
Augmented Aircraft; Wood, J. R. and J. Hodgkinson; 19 December 1980
NLR-TR-79127U
Determination of Low-Speed Longitudinal Maneuvering Criteria for
Transport Aircraft with Advanced Flight Control Systems; Mooij, H. A., W.
P. Boer, and M. F. C. van Gool; 1979
73
MIL-STD-1797A
APPENDIX A
NLR Memorandum
VS-77-024
A Digital Turbulence Model for the NLR Moving - Base Flight Simulator,
Part I; Jansen, C. J., August 1977
NLR Memorandum
VS-77-025
A Digital Turbulence Model for the NLR Moving - Base Flight Simulator,
Part II; Jansen, C. J.; August 1977
Boeing D6-10725
A Simulator and Flight Evaluation of the Longitudinal and Lateral Control
Requirements of the C-5A for the Landing Approach Task; Eldridge, W 18
May 1965
Boeing D6-10732 T/N
A Note on Longitudinal Control Response; Higgins, H. C.; June 1965
Calspan FRM No. 554
The Ideal Controlled Element for Real Airplanes Is Not K/s; Chalk, C. R.-,
August 1981
Comell Aero Lab
IH-2154-F-1
Flight Evaluation of a Stability Augmentation System for Light Airplanes;
Eckhart, F. F., G. W. Hall, and P. A. Martino; November 1966
Cornell Aero Lab
TB-574-F-3
A Flight Investigation of Minimum Acceptable Lateral Dynamic Stability,
Graham, D. and C. James; 30 April 1950
Cornell Aero Lab
TB-574-F-6
A Flight Investigation of Acceptable Roll to Yaw Ratio of the Dutch Roll and
Acceptable Spiral Divergence; Bull, G.; February 1952
Cornell Aero Lab
TB-1094-F-I
Flight Evaluations of the Effect of Variable Spiral Damping in a JTB-26B
Airplane; Rhoads, D. W.; October 1957
Cornell Aero Lab
TB-1444-F-I
Handling Qualities Requirements as Influenced by Prior Evaluation Time
and Sample Size; Kidd, E. A. and G. Bull; February 1963
Douglas Aircraft Co.
LB-25452
Investigation of Pilot-Induced Longitudinal Oscillation in the Douglas Model
A4D-2 Airplane; Terrill, W. H., J. G. Wong, and L. R. Springer; 15 May 1959
General Dynamics Rpt
FZM-12-2652
Norair Rpt No.
NOR-64-143
9 December 1968
Systems Tech. Inc.
TR-124-1
A Systems Analysis of Longitudinal Piloted Control in Carrier Approach,
Cromwell, C. J. and I. L. Ashkenas; June 1962
Systems Tech. Inc.
TR-137-2
Systems Tech. Inc.
TR-189-1
Carrier Landing Analyses; Durand, Tulvio; February 1967
Pilot Induced Oscillations- Their Cause and Analysis; Ashkenas, Irving L.,
Henry R. Jex, and Duane T. McRuer; June 1964
Background Data and Recommended Revisions for MIL-F-8785B(ASG),
'Military Specification -- Flying Qualities of Piloted Airplanes'; Craig, Samuel
J. and Irving L. Ashkenas; March 1971
Systems Tech. Inc.
TR-190-1
Outsmarting MIL-F-8785B(ASG), the Military Flying Qualities Specification;
Stapleford, Robert L., Duane T. McRuer, Roger H. Hoh, et al.; August 1971
Systems Tech. Inc.
TR-199-1
Analytical Assessment of the F-14 Aircraft Control and Handling
Characteristics; Johnston, Donald E. and Samuel J. Craig; February 1972
74
MIL-STD-1797A
APPENDIX A
Systems Tech. Inc.
TR-1090-1
Analytical Assessment of the F-18A Flying Qualities During Carrier
Approach; Ringland, R. F. and D. E. Johnston; September 1977
Systems Tech. Inc.
WP-189-3
Effect of Sideslip on Precise Lateral Tracking; Hoh, R. H. and H. R. Jex;
November 1969
Vought Corp Rpt No.
2-55800/8R-3500
Mathematical Models for the Aircraft Operational Environment of DD-963
Class Ships; Fortenbaugh, R. L.; September 1978
AGARD Rpt 122
The Influence of Drag Characteristics on the Choice of Landing Approach
Speeds; Lean, D. and R. Eaton; 1957
AGARD Rpt 357
Some Low-Speed Problems of High-Speed Aircraft; Spence, A. and D.
Lean, 1961
AGARD Rpt 372
Theory of the Flight of Airplanes in Isotropic Turbulence - Review and
Extension; Etkin, B.; April 1961
AGARD Rpt 420
Flight Measurements of the Influence of Speed Stability on the Landing
Approach; Staples, K. J.; 1963
AGARD-AR-82
The Effects of Buffeting and Other Transonic Phenomena on Maneuvering
Combat Aircraft; Hamilton, B. 1. L.; July 1975
AGARD-AR- 134
Technical Evaluation Report on the Flight Mechanics Panel Symposium on
Stability and Control; Chalk, C. R.; January 1979
AGARD-CP-17
AGARD Stability and Control Meeting, September 1966
“Flying Qualities Criteria Problems and Some Proposed Solutions,"
Carlson, John W. and Richard K. Wilson
“Pilot-Induced Instability;" A'Harrah, R. C. and R. F. Siewert
AGARD-CP- 1 19
Stability and Control; "Flight Simulation - A significant Aid In Aircraft
Design;' A'Harrah, R. C.; April 1972
AGARD-CP-199
Stall/Spin Problems in Military Aircraft; June 1976
AGARD-CP-235
Dynamic Stability Parameters; 'Aircraft Stability Characteristics at High
Angle of Attack;" Kalviste, J.; November 1978
AGARD-CP-249
Piloted Aircraft Environment Simulation Techniques, "Handling Qualities of
a Simulated STOL Aircraft in Natural and Computer-Generated Turbulence
and Shear;" Sinclair, S. R. M. and T. C. West; October 1978
75
MIL-STD-1797A
APPENDIX A
AGARD-CP-260
Proceedings of AGARD Flight Mechanics Panel Symposium on Stability
and Control, September 1978
"Are Today's Specifications Appropriate for Tomorrow's Airplanes?"
A'Harrah, R. C., J. Hodgkinson, and W. J. LaManna
“Flying Qualities and the Fly-by-Wire Aeroplane;" Gibson, J. C.
"L-1011 Active Controls Design Philosophy and Experience;" Urie, David
M.
AGARD-CP-319
Combat Aircraft Maneuverability-, "The Military Flying Qualities
Specification, a Help or a Hindrance to Good Fighter Design?" A'Harrah,
Ralph C. and Robert J. Woodcock; December 1981
AGARD-CP-333
Criteria for Handling Qualities in Military Aircraft; "Simulation for Predicting
Flying Qualities;" Reynolds, P. A.; June 1982
NATO Rpt 408A
Recommendations for V/STOL Handling Qualities; October 1964
NRC of Canada Rpt
LTR-FR- 12
A Flight Investigation of Lateral-Directional Handling Qualities of V/STOL
Aircraft in Low Speed Maneuvering Flight; Doetsch, K. H., et al ; 15 August
1969
RAE Aero. 2504
Problems of Longitudinal Stability Below Minimum Drag, Speed, and
Theory of Stability Under Constraint; Neumark, S.; 1953
RAE Aero. 2688
A Review of Recent Handling Qualities Research, and Its Application to the
Handling Problems of Large Aircraft. Part I: Observations on Handling
Problems and Their Study. Part II: Lateral-Directional Handling; Bisgood,
P. L.; June 1964
RAE TM-FS-46
Developments in the Simulation of Atmospheric Turbulence; Tomlinson, B.
N.; September 1975
RAE TR-68140
Control Characteristics of Aircraft Employing Direct Lift Control, Pinsker, W.
J. G.; May 1968
RAE TR-71021
Glide Path Stability of an Aircraft Under Speed Constraint, Pinsker, \\'. J G.;
February 1971
TSS Standard 5
Supersonic Transport Aeroplane Flying Qualities; 22 May 1964
Ad Hoc Committee Report on B-58 Controllability in Flight, Wright Air Development Division, WrightPatterson AFB, OH, 2 April - 10 May, 1960
Anderson, Ronald O., A Second Analysis of B-58 Flight Control System Reliability, Flight Control
Laboratory, Wright-Patterson AFB, OH, 6 November 1962
Ashkenas, I. L. and T. Durand, "Simulator and Analytical Studies of Fundamental Longitudinal Control
Problems in Carrier Approach,” presented at AIAA Simulation for Aerospace Flight Conference, August.
1963
Behel, I. M. and W. B. McNamara, “F/A-18A High Angle of Attack/Spin Testing," 25th International Report
to the Aerospace Profession, Society of Experimental Test Pilots, September, 1981
76
MIL-STD-1797A
APPENDIX A
Bureau of Naval Weapons Failure Rate Data Handbook, prepared by U. S. Naval Ordnance Laboratory;
Corona, CA (updated periodically)
Caravello, Christopher, Randal G. Joslin, Giuseppe Fristachi, Charles R. Bisbee, Steven S.
Weatherspoon, and Steven G. Henrich, Limited Flight Evaluation as a Function of Aircraft Longitudinal
Dynamics, Air Force Test Pilot School, Class 79A Final Report, December, 1979
Curry, R. E. and A. G. Sim, Unique Flight Characteristics of the AD-1 Oblique-Wing Research Airplane, J
Aircraft, v. 20, nr. 6, June, 1983
"Development of the F/A-18 Handling Qualities Using Digital Flight Control Technology," Society of
Experimental Test Pilots 1982 Report to the Aerospace Profession, 26th Annual Proceedings,
September, 1982
Dryden, Hugh L., “A Review of the Statistical Theory of Turbulence,” Turbulence - Classic Papers on
Statistical Theory, New York: Interscience Publishers, Inc., 1961
Etkin, B., "A Theory of the Response of Airplanes to Random Atmospheric Turbulence," J. Aero/Space
Sciences, July, 1959, 409-420
Etkin, Bernand, Dynamics of Atmospheric Flight, New York: Wiley, 1972
Etkin, Bernard, Dynamics of Flight, New York: Wiley, 1959
Finberg, Floyd, Report of the T-38 Flight Control System PIO Review Board. USAF ASD, February, 1963
Hirsch, Darrell, "Investigation and Elimination of PIO Tendencies in the Northrop T-38A," SAE Paper,
New York, July, 1964
Hodgkinson, J., "Equivalent Systems Approach for Flying Qualities Specification," presented at SAE
Aerospace Control and Guidance Systems Committee Meeting, Denver, CO, 7-9 March, 1979
Hodgkinson, J., R. L. Berger, and R. L. Bear, “Analysis of High Order Aircraft/Flight Control System
Dynamics Using an Equivalent System Approach," presented at 7th Annual Pittsburgh Conference on
Modeling and Simulation, 26-27 April, 1976
Hodgkinson, J., W. J. LaManna, and J. L. Heyde, "Handling Qualities of Aircraft with Stability and Control
Augmentation Systems - A Fundamental Approach," J. R. Ae. S., February, 1976
Houbolt, John C., "Atmospheric Turbulence," AIAA J., Vol. II, No. 4, April, 1973, 421-437
"Industry Observer,” Aviation Week and Space Technology, 1 April, 1968, 13
Jacobson, Ira D. and Dinesh S. Joshi, “Investigation of the Influence of Simulated Turbulence on
Handling Qualities," J. Aircraft, Vol.14, No. 3, March 1977, 272-275
Jones, J. G., "Modeling of Gusts and Wind Shear for Aircraft Assessment and Certification,” Royal Aircraft
Establishment, Paper prepared for CAARC Symposium on Operational Problems, India, October, 1976
Lappe, V. Oscar and Ben Davidson, “On the Range of Validity of Taylor's Hypothesis and the Kilmogoroff
Spectral Law," J. Atmos. Sciences, Vol. 20, November, 1963
Lappe, V. Oscar, “Low-Altitude Turbulence Model for Estimating Gust Loads on Aircraft," J. Aircraft, Vol.
3, No. 1, Jan - Feb, 1966
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Lumley, John L. and Hans A. Panofsky, The Structure of Atmospheric Turbulence, New York:
Interscience Publishers, Inc., 1964
McRuer, Duane, Irving Ashkenas, and Dunstan Graham, Aircraft Dynamics and Automatic Control,
Princeton University Press, 1973
Mitchell, David G. and Roger H. Hoh, “Low-Order Approaches to High-Order Systems: Problems and
Promises," J. Guidance, Control, and Dynamics, Vol. 5, No. 5, Sept - Oct 1982, 482-489
Morgan, Clifford T., Jesse S. Cook, Alphonse Chapanis, and Max W. Lund, eds., Human Engineering
Guide to Equipment Design, New York: McGraw-Hill, 1963
Morgan, Len, “Out for a Spin," Flying, February, 1982
Neal, T. Peter, "Influence of Bobweights on Pilot-Induced Oscillations," J. Aircraft, September, 1971
Otnes, R. K. and L. Enochson; Applied Time Series Analysis, Vol. 1. Basic Techniques; New York- WileyInterscience; 1978
Perkins, Courtland D. and Robert E. Hage, Airplane Performance Stability and Control, New York- Wiley
1949
"Proposals for Revising Mil-F-8785B, 'Flying Qualities of Piloted Airplanes'," AFFDI,-FGC Working, Paper,
February, 1978
Rediess, H. A., D. L. Mallick, and D. T. Berry, Recent Flight Test Results on Minimum Longitudinal
Handling Qualities for Transport Aircraft, presented at the FAUST VIII Meeting, Washington, D.C.,
January 1981
Richards, D. and C. D. Pilcher, "F/A-18A Initial Sea Trials," SETP Cockpit, April/May/June, 1980
Sammonds, R. I., W. E. McNeill, and J. W. Bunnell, "Criteria for Side-Force Control in Air-to-Ground
Target Acquisition and Tracking," J. Aircraft, v. 19, nr. 9, September, 1982
Scott, W. B., “Reengined KC-135 Shows Performance Gains in Test," Aviation Week & Space
Technology, v. 118, nr. 8, McGraw-Hill, February 21, 1983
Stengel, R. F. and G. E. Miller, “Pilot Opinions of Sampling Effects in Lateral-Directional Control,"
presented at 16th Annual Conference on Manual Control, Cambridge, MA, May, 1980
Tentative Airworthiness Objectives and Standards for Supersonic Transport Design Proposals, Flight
Standards Service, FAA, 15 August, 1963
Van Patten, Robert E., Investigation of the Effects of gy. and gz on AFTI/F-16 Control Inputs. Restraints
and Tracking Performance, Interim USAF AMRL Technical Report, August, 1981
von Karman, Theodore, “Progress in the Statistical Theory of Turbulence," Turbulence - Classic Papers
on Statistical Theory, New York: Interscience Publishers, Inc., 1961
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APPENDIX A
30. DEFINITIONS
3.1 Aircraft classification and operational missions.
For the purpose of this standard, the aircraft specified in this requirement is to accomplish the following
missions: _________. The aircraft thus specified will be a Class ____ aircraft. The letter -L following a
class designation identifies an aircraft as land-based; carrier-based aircraft are similarly identified by -C.
When no such differentiation is made in a requirement, the requirement applies to both land-based and
carrier-based aircraft.
REQUIREMENT RATIONALE (3.1)
The very reason for procuring aircraft is to perform one or more missions. The class designation is used
in the handbook to help particularize the requirements according to broad categories of intended use.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 1.3, 1.3.1 and 3.1.1.
Missions
The standard needs a specific mission statement to furnish guidance for interpreting qualitative
requirements as well as for consistent selection of quantitative requirements. Unfortunately, the word
“mission" is used in several contexts not only in this standard, but throughout the writings pertinent to
acquiring a new weapon system. In the broadest sense, ”operational missions” applies to classifying the
aircraft as fighter, bomber, reconnaissance, etc., or to "accomplishing the mission" of bombing, strafing,
etc. In 3.1 the object is to introduce to the designer in general terms the function of the vehicle he is to
design. It should be sufficient for the procuring activity to refer to those paragraphs of the System
Specification and Air Vehicle Specification to define the overall performance requirements, the
operational requirements, employment and deployment requirements.
The operational missions considered should not be based on just the design mission profiles. However,
such profiles serve as a starting point for determining variations that might normally be expected in
service, encompassing ranges of useful load, flight time, combat speed and altitude, in-flight refueling,
etc., to define the entire spectrum of intended operational use. “Operational missions" include training and
ferry missions.
The intended use of an aircraft must be known before the required configurations, loadings, and the
Operational Flight Envelopes can be defined and the design of the aircraft to meet the requirements of
this standard undertaken. If additional missions are foreseen at the time the detail specification is
prepared, it is the responsibility of the procuring activity to define the operational requirements to include
these missions. Examples of missions or capabilities that have been added later are in-flight refueling
(tanker or receiver), aerial pickup and delivery, low-altitude penetration and weapon delivery, and ground
attack for an air-superiority fighter or vice versa.
Once the intended uses or operational missions are defined, a Flight Phase analysis of each mission
must be conducted. With the Flight Phases established, the configurations and loading states which will
exist during each Phase can be defined. After the configuration and loading states have been defined for
a given Flight Phase, Service and Permissible Flight Envelopes can be determined and Operational Flight
Envelopes more fully defined.
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APPENDIX A
For military rotorcraft STI has proposed (NASA CR 177331 or NASA CR 177304) a more detailed
alternative to general category, class and flight phase definitions. Each Flight Phase is assigned specific
appropriate tasks, 3 to 14 in number. Each of these tasks is quantified in terms of a detailed maneuver,
including tolerances on performance. This structure should provide an adequate basis for evaluating
mission-task performance and pilot workload, which are the essence of flying qualities, directly rather
than through the response parameters by which flying qualities are usually specified. Although more
difficult to relate back to design, this alternative provides an excellent set of criteria for assessing
operational worth of the actual vehicle in flight.
Classification of Aircraft--An aircraft is placed in one of the following Classes:
Class I:
Small light aircraft such as:
Light utility
Primary trainer
Light observation
Class II:
Medium weight, low-to-medium maneuverability aircraft such as:
Heavy utility/search and rescue
Light or medium transport/cargo/tanker
Early warning/electronic countermeasures/airborne command,
control, or communications relay
Antisubmarine
Assault transport
Reconnaissance
Tactical bomber
Heavy attack
Trainer for Class II
Class III:
Large, heavy, low-to-medium maneuverability aircraft such as:
Heavy transport/cargo/tanker
Heavy bomber
Patrol/early warning/electronic countermeasures/airborne command,
control, or communications relay
Trainer for Class III
Class IV:
High-maneuverability aircraft such as:
Fighter-interceptor
Attack
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APPENDIX A
Tactical reconnaissance
Observation
Trainer for Class IV
The Class designation aids in selecting and interpreting handbook material. The procuring activity will
assign an aircraft to one of these Classes, and the handbook requirements for that Class are meant to
apply. When no Class is specified in the requirement, the requirement is meant to apply to all Classes.
When operational missions so dictate, an aircraft of one Class should be required by the procuring activity
to meet selected requirements ordinarily specified for aircraft of another Class. The classification scheme
simplifies mission definition. Basically, the four Classes are related qualitatively to maximum design gross
weight and symmetrical flight limit load factor at the basic flight design gross weight, as shown on figure
3.
The presentation of figure 3 makes it obvious that highly maneuverable aircraft such as fighter and attack
types, together with certain trainer and observation craft, should be designed for high limit load factor.
These vehicles tend to group in the weight range from 5000 to 100,000 lb. There are a few small,
lightweight trainers and observation aircraft which are also designed for fairly high load factors, which
could be in either Class I or Class IV. Classification of these aircraft should be on the basis of more
detailed information about the intended use; or alternatively the detail specification should be a
combination of appropriate requirements.
Figure 3 also illustrates that all other aircraft are required to be designed for a limit load factor of less than
4 g, and that current aircraft span the weight range from 1000 to almost 1,000,000 lb. In addition, there
may be significant differences in the way each vehicle responds to atmospheric turbulence or wind shear.
Another factor of possible significance is the location of the pilot in the vehicle relative to the center of
gravity and the extremities of the vehicle. The location of the pilot in the vehicle affects his motions and
ride qualities. If the effects of each of these factors on handling or flying qualities were fully understood
and a sufficient data base existed, then the quantitative requirements could be stated as mathematical or
empirical functions of the significant factors, and there would be no need for any classification breakdown
to accommodate these effects in the specification requirements.
It should also be recognized that as vehicles become larger, practical design considerations may dictate
compromises between the degree of maneuverability and the values of flying qualities parameters that
are desirable and what can be accepted, through relaxation of operational requirements or through
modification of operational procedures or techniques.
How best to handle the factors discussed above is not completely clear at this time. Ideally the
requirements should be expressed as mathematical functions of the significant factors. The current state
of knowledge and the experimental data available do not permit this, so it is necessary to make the
relatively arbitrary Class definition. Further research into possible scaling parameters, simulation study,
and operational experience is required in this area.
REQUIREMENT LESSONS LEARNED
In keeping with overall guidance to relate requirements to the intended mission, this mission statement
has been found to be necessary to guide selection of flight conditions and tailoring of requirements.
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MIL-STD-1797A
APPENDIX A
FIGURE 3. Classification of aircraft (AFFDL-TR-69-72).
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MIL-STD-1797A
APPENDIX A
3.2 Flight Phase Categories.
To accomplish the mission requirements the following general Flight Phase Categories are involved:
_____________. Special Flight Phases to be considered are: _____________.
REQUIREMENT RATIONALE (3.2)
Flying qualities requirements vary for the different phases of a mission. To the extent permitted by
available data, these variations have been taken into account.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 1.4.
The appropriate Flight Phases to be listed are to be chosen from the material that follows. Experience
with aircraft operations indicates that certain Flight Phases require more stringent values of flying qualities
parameters than do others (e.g., air-to-air combat requires more dutch roll damping than does cruising
flight). Also, a given mission Flight Phase will generally have an Aircraft Normal State associated with it
(e.g., flaps and gear down for landing approach and up for cruising flight; maximum gross weight).
In flight and simulator evaluations, pilots generally rate a set of flying qualities on suitability for a given
mission segment like one of these Flight Phases. The pilots assign an overall rating, based on ability and
effort required to perform certain appropriate tasks such as precision tracking of a target or a glide slope,
trimming and making heading changes at constant altitude, in an appropriate environment. The similarity,
of tasks in many Flight Phases, plus the limited amount of evaluation data on specific Flight Phases, has
led to grouping the Phases into three Categories.
Nonterminal Flight Phases:
Category A.
These nonterminal Flight Phases that require maneuvering, precision
tracking, or precise flight-path control. Included in this Category are:
a. Air-to-air combat (CO)
b. Ground attack (GA)
c. Weapon delivery/launch (WD)
d. Aerial recovery (AR)
e. Reconnaissance (RC)
f. In-flight refueling (receiver) (RR)
g. Terrain following (RF)
h. Antisubmarine search (AS)
i. Close formation flying (FF)
j. Low-altitude parachute extraction (LAPES) delivery
Category B:
Those nonterminal Flight Phases that are normally accomplished using
gradual maneuvers and without precision tracking, although accurate
flight-path control may be required. Included in this Category are:
a. Climb (CL)
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APPENDIX A
b. Cruise (CR)
c. Loiter (LO)
d. In-flight refueling (tanker) RT)
e. Descent (D)
f. Emergency descent (ED)
g. Emergency deceleration (DE)
h. Aerial delivery (AD)
Terminal Flight Phases:
Category C:
Terminal Flight Phases are normally accomplished using gradual
maneuvers and usually require accurate flight-path control. Included in
this Category are:
a. Takeoff (TO)
b. Catapult takeoff (CT)
c. Approach (PA)
d. Waveoff/go-around (WO)
e. Landing (L)
When necessary, the procuring activity may specify recategorization or addition of Flight Phases or
delineation of requirements for special situations, e.g., zoom climbs.
These Flight Phases are to be considered in the context of the total mission so that there will be no gap
between successive Phases of any flight, and so that transition will be smooth. In certain cases,
requirements are directed at specific Flight Phases identified in the requirement. When no Flight Phase or
Category is stated in a requirement, that requirement is meant to apply to all three Categories.
For the most part, the Flight Phase titles are descriptive enough to facilitate picking those applicable to a
given design. The Formation Flying (FF) Flight Phase is intended to be used, if desired, where there is no
other requirement for rapid maneuvering, precision tracking, or precise flight-path control in up-and-away
flight. An example might be a Class I trainer for which the procuring activity desires Category A flying
qualities (note the use of the T-37, T-38, etc. in non-training roles).
Not all of these Flight Phases apply to a given aircraft. Those that are appropriate to design operational
missions and emergencies will be chosen for each design. The list cannot be exhaustive because new
mission requirements continue to be generated. Thus the procuring activity may delete some Phases and
add others. Responsibility for choosing applicable Flight Phases, as with filling in most or all of the blanks,
is initially the procuring activity's. The contractor should assure that this listing is inclusive and exhaustive
(for the stated primary and alternate missions), and suggest necessary additions. It is the procuring
activity's responsibility either to agree with the contractor's suggestions or to recategorize the Flight
Phases.
In certain cases, both flying qualities requirements and aircraft capabilities may be less than one would
ordinarily expect. An example is a zoom climb--a dynamic maneuver in which qualities such as speed
stability and natural frequency cannot be measured in flight, and the effectiveness of aerodynamic
controls
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MIL-STD-1797A
APPENDIX A
is necessarily low at low dynamic pressure. Lacking enough data to formulate general quantitative
requirements for these cases, we leave for the procuring activity the provision of specific requirements as
specific mission needs dictate.
For each Flight Phase or Flight Phase Category (depending upon the data available) typical flight
conditions, maneuvers, disturbances, side tasks, etc. have been assumed in setting the suggested
numerical values. The accurate flight-path control for landing, as an example, may well be a high-pin
piloting task to which some Category A requirements apply. In tailoring the requirements for a particular
procurement, any envisioned operating conditions more lax or more stringent than normal should be
taken into account to the extent possible.
REQUIREMENT LESSONS LEARNED
As an example of the last caveat above, consider the A-10 experience documented by Brandeau in
AFFDL-TR-78-171. That airplane appeared to meet MIL-F-8785B Level 1 requirements for Category A
(which includes ground attack) and it was rated Level 1 during flight tests using a straight-in approach. Its
flying qualities were unsatisfactory, however, when evaluated in an operationally realistic ground attack
task.
In close air support, a wide variety of attack maneuvers may be characterized by three general phases,
as shown in figure 4:
FIGURE 4. Ground attack maneuver scenario.
Target acquisition - Rapid rolling toward target while developing 4 to 5 g's; bank and g's held until rollout
onto target (return to zero bank and 1 g)
Weapon delivery or tracking/firing - errors eliminated and pipper maintained on target
Break - a gross maneuver to reposition for another attack while looking after aircraft survival.
For gross target acquisition maneuvers, highly predictable terminal orientation of the velocity vector is
vital in order to minimize the duration of the relatively vulnerable weapon delivery phase. Excellent roll
response is required, in terms of both quickness and maintaining turn coordination. Weapon delivery
requires rapid, precise control of the velocity vector for dropping unguided bombs, or of the pipper line of
sight (and thus aircraft attitude) for gunnery.
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MIL-STD-1797A
APPENDIX A
While the original A-10 stability augmentation apparently met MIL-F-8785B requirements on lateraldirectional dynamics, and pilots rated it satisfactory in “the originally planned tactical maneuvers... It was
only as the maneuvers became very aggressive that the problem surfaced.” For these aggressive
maneuvers, the average maxima quoted are:
normal acceleration
4.5 g
roll rate
93 deg/sec
bank angle
93 deg
tracking time
2.33 sec
To satisfy the requirements of the task outlined above, the aerodynamic configuration remained
unchanged and the flight control system modifications were relatively minor. This will not necessarily be
so in more sophisticated designs. The cost of fixing such deficiencies could be very high after a new
aircraft has flown, and so it would obviously be beneficial to consider operational maneuvers as early as
possible in the design phase. In the example cited, little more than figure 4 would be required as an
additional Flight Phase in the specification.
For this more severe Flight Phase, more stringent requirements might be placed on Dutch roll damping
and roll-yaw coupling--see figure 5 responses of lateral tracking error to a roll control doublet. Although
the A-10 deficiency was indicated at high g's, certainly for such a severe Flight Phase the lateraldirectional characteristics must be investigated in pullups and turns - and roll-sideslip coupling in rapid
rolls-as well as in straight flight. (While the requirements of MIL-F-8785B apply throughout the V-h-n Flight
Envelopes, often the lateral-directional behavior has been evaluated primarily in 1-g flight.) Commonly it
is observed that the amount of aileron-to rudder crossfeed needed to coordinate turn entries varies
considerably with angle of attack. Thus, one might find no single crossfeed gain suitable for all phases of
the ground attack described.
FIGURE 5. Response to a 1-second, half-stick aileron doublet for tracking scenario.
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MIL-STD-1797A
APPENDIX A
In addition, advancing flight control technology has greatly increased the potential for tailoring the flying
qualities for specific tasks within a Flight Phase Category without compromising other tasks. Truly taskoriented flying qualities would receive an impetus from the inclusion of requirements related to actual
operational tasks into the specification for a particular aircraft.
An example of the need for better flight characteristics for an added task is the low-altitude parachute
extraction mission, in which the pilot must fly precisely at very low altitude. According to an Air Force test
pilot, for this task the C-130 is “Level 2 at best, mostly Level 3”. The lesson again is to account for the
flying qualities implications of changes in operational usage.
To the extent feasible, we have tailored the requirements to particular tasks of the Flight Phases. A very
important, but unstated, corollary is the need to avoid inconsistencies in flight control mechanization from
one Flight Phase or configuration to another. Drastic or numerous changes in control mode have the
potential to confuse the pilot, to the detriment of mission effectiveness or even flight safety. With few
exceptions, a single flight technique should suit all operations.
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MIL-STD-1797A
APPENDIX A
3.3 Levels and qualitative suitability of flying qualities.
The handling characteristics described in this standard are specified in terms of qualitative degrees of
suitability and Levels. The degrees of suitability are defined as:
Satisfactory
Flying qualities clearly adequate for the mission Flight Phase. Desired
performance is achievable with no more than minimal pilot compensation
Acceptable
Flying qualities adequate to accomplish the mission Flight Phase, but some
increase in pilot workload or degradation in mission effectiveness, or both,
exists
Controllable
Flying qualities such that the aircraft can be controlled in the context of the
mission Flight Phase, even though pilot workload is excessive or mission
effectiveness is inadequate, or both. The pilot can transition from Category A
Flight Phase tasks to Category B or C Flight Phases, and Category B and C
Flight Phase tasks can be completed.
Level 1 is Satisfactory, Level 2 is Acceptable, and Level 3 is Controllable. In the presence of higher
intensities of atmospheric disturbances, 4.9.1 states the relationship between Levels and qualitative
degrees of suitability. Where possible, the flying qualities requirements are stated for each Level in terms
of limiting values of one or more parameters. Each value, or combination of values, represents a
minimum condition necessary to meet one of the three Levels of acceptability.
It is to be noted that Level 3 is not necessarily defined as safe. This is consistent with the Cooper-Harper
rating scale: for Cooper-Harper ratings of 8 and 9, controllability may be in question. If safe characteristics
are required for Level 3, then action must be taken to improve aircraft flying qualities.
In some cases sufficient data do not exist to allow the specification of numerical values of a flying quality
parameter. In such cases it is not possible to explicitly define a quantitative boundary of each Level, so
the required Levels are then to be interpreted in terms of qualitative degrees of suitability for the piloting
task, appropriate for mission accomplishment.
REQUIREMENT RATIONALE (3.3)
These Levels and degrees of suitability are part of the structure of the standard, and are based on the
Cooper-Harper Scale, see figure 6 (NASA-TN-D-5153).
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 1.5.
Where possible, the requirements of Section 4 have been stated as allowable ranges of the stability or
control parameter being specified. Each specified value is a minimum condition to meet one of three
Levels of acceptability related to the ability to complete the operational missions for which the aircraft is
designed. In actual practice, the flying quality boundaries referred to above were obtained by fairing lines
of constant Cooper-Harper pilot rating. Hence it was necessary to define equivalent definitions between
the Cooper-Harper scale shown in figure 6 and the Level definitions. Typically, a Cooper-Harper pilot
rating of 1 to 3 defines Level 1, a Cooper-Harper rating from 4 through 6 defines Level 2, and a CooperHarper rating from 7 through 9 defines Level 3. In a few instances, as indicated in the associated
Guidance, the boundaries have been further modified for other considerations. It should be noted that the
Level 3 Controllable lower bound, like the other bounds, is sensitive to pilot workload. AFFDL-TR-69-72
shows the assumed correlation of the Cooper-Harper scale with earlier rating scales.
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MIL-STD-1797A
APPENDIX A
FIGURE 6. Definition of Flying Quality Levels in Calm to Light Turbulence.
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MIL-STD-1797A
APPENDIX A
Fractional ratings are to be avoided: Cooper and Harper stress the need for clear choices, especially at
the Level boundaries. Wildly scattered ratings should not be averaged: the Cooper-Harper scale is
nonlinear. Even for adjacent ratings, caution is needed: variations in pilot technique, disturbance time
histories, subjective criteria, etc. may result in valid rating differences.
Relating the Cooper-Harper Pilot Rating Scale to the Levels of flying qualities has the added benefit of
more precise definitions which are related to the operational considerations of pilot workload and task
performance, as well as making the pilot rating correlations consistent with the Level 1, 2, and 3 criterion
boundaries in the flying quality standard. It is especially important to note that "Controllable" is in the
context of the Flight Phase: the pilot's other duties must be attended to.
It is natural for pilot rating of flying qualities to degrade with increasing atmospheric disturbances. Since
this standard is used to procure aircraft, not pilots, we must distinguish between degradation of pilot rating
and degradation of aircraft characteristics. As indicated in the requirement, this distinction is made in
4.9.1 for Normal and Failure States. These allowances, of course, should not be construed as a
recommendation to degrade flying qualities with increasing intensities of atmospheric disturbances.
For several reasons we do not use Cooper-Harper ratings directly in the standard:
Level applies to aircraft (which the requirements cover), doing design-mission task. CooperHarper (C-H) rating is given by a pilot doing the task with the aircraft in a given environment.
Since C-H rating is expected to change with severity of the environment, Levels tied exclusively
to C-H ratings would vary with intensity of disturbances--e.g. Level 1 in light turbulence, Level 2 in
moderate. That gets cumbersome to call out in requirements, so we need to tie down the
environmental severity when determining Levels.
Requirements need to address ability to complete or terminate a Flight Phase, which the C-H
ratings don't treat.
We need some leeway for engineering-type input to requirements, e.g. increasing the Level 1
short-period damping boundary to account for more severe turbulence and not allowing negative
dutch roll damping even for Level 3.
Some requirements are based on operational experience or need, rather than pilot evaluation-e.g., crosswind landing capability or two-engine-out controllability.
"Deficiencies warrant improvement,” from the C-H ratings, just doesn't apply for normal operation
between the Operational and Service Flight Envelope boundaries.
In assessing compliance, as well as in design, the firmness of numerical values is preferred to the
variability of evaluation-pilot ratings. Thus, even though the final decision to accept or reject will
be made on the basis of pilot ratings (as these decisions always have been), more exactly
defined boundaries are needed for specification.
Having stated these caveats, we note that given a well-defined task and a "calm to light" environment, the
Level definitions do closely correspond to 1 through 3, 4 through 6, and 7 through 9 C-H ratings. Thus, for
the qualitative requirements in the proper environment there is a direct correspondence of C-H ratings, to
Levels (However it's done, lacking the "design" environment the evaluation pilot must extrapolate). We
have modified the Level definitions to be even closer to the C-H definitions.
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MIL-STD-1797A
APPENDIX A
REQUIREMENT LESSONS LEARNED
Use of the Cooper-Harper scale is accepted universally as a guide, but only by some as a way to state
requirements. Herein we use it as the principal way to relate flying qualities requirements to operational
needs. Accounting for the observed effects of atmospheric disturbances in a generally acceptable manner
has been quite a problem. The point is that while pilot rating is allowed to degrade in Moderate
disturbances, as we must expect, we do not want to allow aircraft characteristics also to degrade, as they
might from saturation of stability augmentation or other nonlinearities. That would likely cause a further
degradation in pilot rating. Therefore we must somehow make a distinction between Levels and CooperHarper ratings; their relationship must vary with the intensity of atmospheric disturbances.
According to figure 6, piloting should not require the pilot's attention to the exclusion of all other duties. If
“adequate performance requires moderate [or greater] pilot compensation", the “deficiencies warrant
improvement", etc.--all in the full context of the mission including hostile environment, weather, etc. The
ratings are meant to apply to the most demanding tasks foreseen, aggressively performed.
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MIL-STD-1797A
APPENDIX A
40. REQUIREMENTS
4.1 General requirements
4.1.1 Loadings.
The contractor shall define the longitudinal, lateral and vertical envelopes of center of gravity and
corresponding weights that will exist for each Flight Phase. Throughout these envelopes shall include the
most forward and aft center-of-gravity positions as defined in ______________. In addition the contractor
shall determine the maximum center-of-gravity excursions attainable through failures in systems or
components, such as fuel sequencing or hung stores, for each Flight Phase. Throughout these
envelopes, plus a growth margin of __________, and for the excursions cited, this standard applies.
REQUIREMENT RATIONALE (4.1.1)
Since aircraft characteristics vary with loading, limits must be defined and the loadings known at
conditions for demonstration of compliance. The loading of an aircraft is determined by what is in (internal
loading) and attached to (external loading) the aircraft. The loading parameters that normally define flying
qualities are weight, center-of-gravity position, and moments and products of inertia (4.1.2). External
stores affect all these parameters and also affect aerodynamic coefficients.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.1.2.
For normal operation the allowable c.g. range is applicable. Certain failures may cause an adverse c.g.
shift. In these cases, the abnormal c.g.'s so attained are applicable.
The requirements apply under all loading conditions associated with an aircraft's operational missions.
Since there are an infinite number of possible internal and external loadings, each requirement generally
is only examined at the critical loading(s) with respect to the requirement. Only permissible center-ofgravity positions need be considered for Aircraft Normal States. Fuel sequencing, transfer failures or
malperformance, and mismanagement that might move the center of gravity outside the established limits
are expressly to be considered as Aircraft Failure States. The worst possible cases that are not approved
Special Failure States (4.1.7.2) must be examined.
MIL-W-25140 is normally referenced here for consistency.
The procuring activity may elect to specify a growth margin in c.g. travel to allow for uncertainties in
weight distribution, stability level and other design factors, and for possible future variations in operational
loading and use. Peculiarities of configuration or possible alternative mission tasks may lead to the
specification of additional loadings. Fuel slosh and shift under acceleration also need consideration.
It is fairly straightforward to determine those longitudinal flying qualities that set the longitudinal c.g. limits,
but there are also cases where the aft c.g. limit may be set by lateral-directional flying qualities. Usually
supersonic flight at high dynamic pressure is the most critical, because the level of directional stability is
reduced due to Mach number and aeroelastic effects. Conditions to be investigated to ensure acceptable
lateral-directional characteristics at the aft limits include:
a. Roll performance/roll coupling
b. Abrupt engine loss or inlet unstart at one g
c. Abrupt engine loss or inlet unstart at high normal acceleration or angle of attack, especially for
Class IV aircraft
d. Turbulence effects
92
MIL-STD-1797A
APPENDIX A
When evaluating aeroelastic effects and effects of hinge moments due to angle of attack, the pitch and
yaw control surface deflection and sideslip must be taken into account. Pitching moment due to sideslip
can also be significant on configurations using vortex lift or highly swept wings. The trend toward relaxed
lateral-directional stability indicates that the effect of c.g. on low-speed lateral-directional characteristics
should also be examined.
This requirement also requires the contractor to define the lateral c.g. limit. This is especially critical for
Class IV aircraft. Conditions to evaluate with asymmetric loadings include:
a. Takeoff with and without crosswind.
b. Roll performance/roll coupling.
c.
Abrupt engine loss at takeoff and in maneuvering flight.
d. Dive pullout at high normal acceleration.
e. Yaw departure at high angle of attack and spin resistance.
In defining this limit, the basic lateral asymmetry due to wing fuel system tolerances and equipment
mounted off centerline, such as guns and ammunition, should be taken into account.
REQUIREMENT LESSONS LEARNED
Lateral asymmetries due to fuel loading can have important effects on trim, stall/post-stall characteristics.
etc. Fuel system design has been known to promote such asymmetry, for example, at prolonged small
sideslip in cruising flight.
Since the requirements apply over the full range of service loadings, effects of fuel slosh and shifting
should be taken into account in design. Balance, controllability, and airframe and structure dynamic
characteristics may be affected. For example, takeoff acceleration has been known to shift the c.g.
embarrassingly far aft Aircraft attitude may also have an effect. Other factors to consider are fuel
sequencing, in-flight refueling if applicable, and all arrangements of variable, disposable and removable
items required for each operational mission.
5.1 General requirements
5.1.1 Loadings-verification.
The contractor shall furnish the required loading data in accordance with the Contract Data Requirements
List (CDRL).
VERIFICATION RATIONALE (5.1.1)
Aircraft weight and balance are estimated during the design and measured on the vehicle itself.
VERIFICATION GUIDANCE
Once the specific loadings are defined, application of this requirement is straightforward. Provision of this
data is usually called out in the CDRL.
The procuring activity will check the material submitted for completeness. Eventually, weight and balance
measurements will be made to confirm the estimates. The requirements apply to the actual flight weights
and centers of gravity.
VERIFICATION LESSONS LEARNED
93
MIL-STD-1797A
APPENDIX A
4.1.2 Moments and products of inertia.
The contractor shall define the moments and products of inertia of the aircraft associated with all loadings
of 4.1.1. The requirements of this standard shall apply for all moments and products of inertia so defined.
REQUIREMENT RATIONALE (4.1.2)
Inertial characteristics of the aircraft affect its flying qualities, so the contractor must define the inertias for
all expected loadings corresponding to possible distributions and the loadings for which flying qualities
are evaluated.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.1.3.
In any dynamic analysis of an aircraft, inertias must be known. The axis system in which the values are
given must be identified.
Ixy and Iyz , normally neglected, may be appreciable for asymmetric loadings.
REQUIREMENT LESSONS LEARNED
5.1.2 Moments and products of inertia-verification.
The contractor shall furnish moments and products of inertia data in accordance with the Contract Data
Requirements List (CDRL).
VERIFICATION RATIONALE (5.1.2)
By meticulous accounting of weight and balance, inertia values can be estimated fairly accurately. Except
for extremely large aircraft, inertias may be measured with equipment located at the Air Force Flight Test
Center, Edwards AFB, CA.
VERIFICATION GUIDANCE
Sufficient data should be supplied in the reports required by the CDRL. The procuring activity may, at its
discretion, wish to review the methods used in estimating or measuring the inertial characteristics
specified. If deemed necessary, checks of estimates can be made by ground tests (e.g., forced
oscillations using equipment such as that at the Air Force Flight Test Center) or parameter estimation
from flight test data.
VERIFICATION LESSONS LEARNED
94
MIL-STD-1797A
APPENDIX A
4.1.3 Internal and external stores.
The symmetric and asymmetric store combinations to be considered are as follows: _______. The
requirements of this standard shall apply to these store conditions. The effects of stores on the weight,
moments of inertia, center-of-gravity position, and aerodynamic characteristics of the aircraft shall be
determined for each mission Flight Phase. When the stores contain expendable loads, the requirements
of this standard apply throughout the range of store loadings, including sloshing/shifting.
REQUIREMENT RATIONALE (4.1.3)
Once the procuring activity has specified the stores to be considered, the contractor must assure that
evaluation of the aircraft with these store combinations covers all operational flight conditions.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.1.4.
Stores and stores combinations affect the overall ability of the aircraft to meet its mission requirements In
determining the range of store loadings to be specified in the contract, the procuring activity should
consider such factors as store mixes, possible points of attachment, and asymmetries--initial, after each
pass, and the result of failure to release. The contractor may find it necessary to propose limitations on
store loading to avoid excessive design penalties. Since such limitations are operationally restrictive, the
procuring activity may be reluctant to approve them.
The designer should attempt to assure that there are no restrictions on store placement on the aircraft
within the range of design stores. However, it is recognized that occasionally this goal will be
impracticable on some designs. It may be impossible to avoid exceeding aircraft limits, or excessive
design penalties may be incurred. Then, insofar as considerations such as standardized stores permit, it
should be made physically impossible to violate necessary store loading restrictions. If this should not be
practicable, the contractor should submit both an analysis of the effects on flying qualities of violating the
restrictions and an estimate of the likelihood that the restrictions will be exceeded.
Even while the aircraft is in service, other stores will be added. Part of this clearance will be determination
of the effect on flying qualities.
See MIL-HDBK-244 for additional requirements and guidance. The most stringent requirement will
govern.
REQUIREMENT LESSONS LEARNED
High-angle-of-attack testing conducted on the F-15 (AFFTC-TR-75-32) shows a degradation in flying
qualities and departure resistance with external stores. Asymmetric stores and internal fuel asymmetry
affected departure and spin characteristics. Stores tests with the F-16 (AFFTC-TR-80-29) show similar
results, and serve to illustrate the importance of defining a comprehensive set of conditions for
investigating stores effects. For example, during ground taxi of the F-16, “The pilot noticed a leaning or
tip-over sensation especially during light weight taxi with a strong crosswind, tight turns, or with
asymmetric store loadings.”
Stores can have mass, inertial and aerodynamic effects, typically decreasing both longitudinal and
directional aerodynamic stability, increasing moments of inertia and the roll-mode time constant, and
increasing susceptibility to departure from controlled flight and the difficulty of recovery. The available
control power limits the amounts of inertia increase and instability that can be tolerated. Store separation
is a prime concern.
For external as well as internal fuel tanks, the critical fuel loading for each requirement applies (see 4.1.1)
95
MIL-STD-1797A
APPENDIX A
5.1.3 Internal and external stores-verification.
The analysis, simulation and testing to verify compliance with this standard shall include the listed stores.
The contractor shall furnish a list of store restrictions in accordance with the Contract Data Requirements
List (CDRL)
VERIFICATION RATIONALE (5.1.3)
The c.g. and aerodynamic effects of internal and external stores need to be considered while verifying
compliance with the flying qualities requirements.
VERIFICATION GUIDANCE
Often the large number of possible stores combinations will, from a practical standpoint, limit flight
demonstration to a few cases. A careful analysis before flight testing will assure that the most critical
combinations (from a flying qualities perspective) are being evaluated.
VERIFICATION LESSONS LEARNED
96
MIL-STD-1797A
APPENDIX A
4.1.4 Flight Envelopes.
Three kinds of envelopes are to be defined. They are the Operational Flight Envelope, the Service Flight
Envelope, and the Permissible Flight Envelope. As a general policy, the contractor should propose the
boundaries and rationale for all envelopes. Additional negotiation between the contractor and flying
qualities engineer representing the procuring agency may be required before the boundaries of the flight
envelopes can be agreed to by the procuring activity.
At this stage the Flight Phases will also be known from 3.2. In response to these and other requirements
the contractor can then design the aircraft and:
Further define the Operational Flight Envelope for each Flight Phase, based on the associated
Aircraft Normal States,
Construct the larger Service Flight Envelope for the Aircraft Normal State associated with each
Flight Phase, and
Similarly construct portions of the Permissible Flight Envelope boundaries, beyond which
operation is not allowed.
Each Envelope must include the flight condition(s) related to any pertinent performance guarantees.
The envelopes are described by the specification of a two-dimensional (speed and load factor) figure
representing the conditions where the requirements apply at a given altitude. An example that defines
terms for the Operational and Service Envelopes is shown on figure 7. The load factor, n, denotes
maneuverability without regard to thrust available, i.e., the flying qualities specification places no
requirements on load-factor capability in constant-speed level flight. These Envelopes are defined at
various altitudes corresponding to the Flight Phases; thus they could be considered to be threedimensional.
FIGURE 7. Definition of Flight Envelope terms.
For a given design, angle of attack usually will be a more succinct bound than speed, altitude and load
factor that vary with gross weight. For relation to aircraft missions the envelopes should be kept in
standard V, h, n form, but nominal weight can be used if it and the limit angle of attack are given. Flight
testing is likely to use angle of attack directly.
97
MIL-STD-1797A
APPENDIX A
Some Flight Phases of the same Category will involve the same, or very similar. Aircraft Normal States;
so one set of flight envelopes may represent several Flight Phases. Each Flight Phase will involve a
range of loadings. Generally, it will be convenient to represent this variation by superimposing boundaries
for the discrete loadings of 4.1.1, or possibly by bands denoting extremes. If different external store
complements affect the envelope boundaries significantly, it may be necessary for the contractor to
construct several sets of envelopes for each Flight Phase, each set representing a family of stores. A
manageably small total number of Envelopes should result. It is apparent that the Flight Envelopes must
and can be refined, as the design is further analyzed and defined, by agreement between the contractor
and the procuring activity.
Flight tests will be conducted to evaluate the aircraft against requirements in known (a priori) Flight
Envelopes. Generally, flight test will cover the Service Flight Envelope, with specific tests (stalls, dives,
etc.) to the Permissible limits. The same test procedures usually apply in both Service and Operational
Envelopes; only the numerical requirements and qualitative Levels differ. If, for example, speed and
altitude are within the Operational Flight Envelope but normal load factor is between the Operational and
Service Flight Envelope boundaries, the requirements for the Service Flight Envelope apply for Aircraft
Normal States.
Ideally, the flight test program should also lead to definition of Flight Envelopes depicting Level 1 and
Level 2 boundaries. These Level boundaries should aid the using commands in tactical employment,
even long after the procurement contract has been closed out.
Separate Flight Envelopes are not normally required for Aircraft Failure States. It is rational to consider
most failures throughout the Flight Envelopes associated with Aircraft Normal States. These may be
exceptions (such as a wing sweep failure that necessitates a wing-aft landing, or a flap failure that
requires a higher landing speed) that are peculiar to a specific design. In such cases the procuring activity
may have to accept some smaller Flight Envelopes for specific Failure States, making sure that these
Envelopes are large enough for safe Level 2 or Level 3 operation.
Level 2 flying qualities are required in the Service Flight Envelope. Note, however, that the minimum
service speed is a function of stall speed, Vs, and the first item in the definition is based on lift plus thrust
component. For STOL or high-thrust-to-weight-ratio configurations, VS by this definition can be
significantly lower than the aerodynamic or power-off stall speed. Other items in the definition of VS and
minimum service speed give a minimum usable speed which could be higher or lower than the
aerodynamic stall speed. This applies in level flight and in maneuvers. It is doubtful that this interpretation
has in fact been used; however, these are operational benefits to be gained from improving flying
qualities at extreme flight conditions. The safe, usable attainment of more extreme flight conditions may
be emphasized for missions in which maneuvering at high angle of attack is critical. The procuring activity
could accomplish this by, tailoring the requirements for determining the Service and Permissible Flight
Envelopes. As an example, we could require that the Permissible Flight Envelope be defined consistent
with operational maneuvers appropriate to the mission.
In the roll performance requirements we have felt the need to make a further distinction as a function of
airspeed within the Operational and Service Flight Envelopes. The relaxation close to the stall is a
concession to aerodynamic realities for roll control via the usual means.
It should also be noted that the boundaries of these envelopes should not be set by ability to meet the
flying qualities requirements. The flying qualities requirement should be met within the boundaries which
normally are set by other factors, unless specific deviations are granted. The only exception is control
power, which may set some boundaries for stable aircraft, if the requirements on the Operational Flight
Envelope are still met. The rationale for each type of Envelope is presented in the following discussions of
each subparagraph.
The Air Force Flight Test Center has expressed the desirability of more flight testing away from the middle
of the flight envelopes, where most of the testing is done.
98
MIL-STD-1797A
APPENDIX A
4.1.4.1 Operational Flight Envelopes.
The Operational Flight Envelopes define the boundaries in terms of speed, altitude and load factor within
which the aircraft must be capable of operating in order to accomplish the missions of 3.1 and in which
Level 1 flying qualities are required. These envelopes shall implicitly include the ranges of other
parameters, such as sideslip, which may normally be encountered. The range of sideslip or lateral
acceleration employed with direct side force control is to be stated explicitly. In the absence of other
specific instructions, the contractor shall use the representative conditions of table I for the applicable
Flight Phases.
REQUIREMENT RATIONALE (4.1.4.1)
By bounding the envelopes in which the best flying qualities are desired, unnecessary cost, weight,
complexity, etc. can be avoided while assuring capability to perform the intended missions.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.7.
Operational Flight Envelopes are regions in speed-altitude-load factor space where it is necessary for an
aircraft, in the configuration and loading associated with a given Flight Phase, to have very good flying
qualities -- as opposed, for example, to regions where it is only necessary to insure that the aircraft can
be controlled without undue concentration. The required size of the Operational Flight Envelopes for a
particular aircraft has been given in 3.1; however, this can further be delineated by using table VII for
each Flight Phase Category. Additional boundaries will be provided by the contractor.
It has not generally been found necessary to incorporate other flight parameters such as sideslip; the
operationally encountered range of such parameters is implicit in the Operational Flight Envelopes. Also,
certain requirements specify a sideslip range or capability.
In defining the speed-altitude-load factor combinations to be encompassed, the following factors should
be considered:
a.
The Operational Flight Envelope for a given Flight Phase should initially be considered to
be as large a portion of the associated Service Flight Envelope as possible, to permit the
greatest freedom of use of the aircraft.
b.
The detail specification should be as specific as possible about the speed and altitude
ranges over which stated load-factor capabilities are required. Obviously, limit load factor
cannot be attained at a lift-limited combat ceiling; but normally it would be insufficient at a
much lower altitude to have nL capability at only one speed.
c.
The Operational Flight Envelopes must encompass the flight conditions at which all
appropriate performance guarantees will be demonstrated.
d.
In setting the minimum approach speed, VOmin(PA), care should be taken to allow
sufficient stall margin. Commonly, 1.2 VS has been used for military land-based aircraft
and 1.15 VS for carrier-based aircraft. For reference, FAR Part 25 specifies 1.3 Vs for
landing-distance calculations; while FAR Part 23 specifies approach at 1.5 VS for these
calculations where required. Differences in method of determining stall speed tend to
make the FAR Part 25 and military stall margins equivalent; herein VS is defined at
constant airspeed rather than with a slow deceleration.
99
MIL-STD-1797A
APPENDIX A
Table I. Operational Flight Envelope.
FLIGHT
PHASE
AIRSPEED
FLIGHT PHASE
CATEGORY
A
B
C
100
ALTITUDE
LOAD FACTOR
MIL-STD-1797A
APPENDIX A
TABLE VII. Operational flight Envelope values.
FLIGHT
PHASE
CATEGORY
A
B
C
AIRSPEED
FLIGHT PHASE
Vo
(Mo
min
)
min
ALTITUDE
Vo
max
(Mo
max
)
LOAD FACTOR
ho
ho
no
min
max
no
min
max
AIR-TO-AIR COMBAT (CO)
1.15VS
VMAT
MSL
Combat
Ceiling
-1.0
nL
GROUND ATTACK (GA)
1.3 VS
VMRT
MSL
Medium
-1.0
nL
WEAPON DELIVERY/LAUNCH
(WD)
Vrange
VMAT
MSL
Combat
Ceiling
.5
*
AERIAL RECOVERY (AR)
1.2 VS
VMRT
MSL
Combat
Ceiling
.5
nL
RECONNAISSANCE (RC)
1.3 Vs
VMAT
MSL
Combat
Ceiling
*
*
IN-FLIGHT REFUEL (RECEIVER)
(RR)
1.2 Vs
VMRT
MSL
Combat
Ceiling
.5
2.0
TERRAIN FOLLOWING (TF)
Vrange
VMAT
MSL
10,000 ft
0
3.5
ANTISUBMARINE SEARCH (AS)
1.2 VS
VMRT
MSL
Medium
0
2.0
CLOSE FORMATION FLYING (FF)
1 4 VS
VMAT
MSL
Combat
Ceiling
-1.0
nL
CLIMB (CL)
.85 VR/C
1.3 VR/C
MSL
Cruising
Ceiling
.5
2.0
CRUISE (CR)
Vrange
VNRT
MSL
Cruising
Ceiling
.5
2.0
LOTTER (LO)
.85 Vend
1.3 Vend
MSL
Cruising
Ceiling
.5
2.0
IN-FLIGHT REFUEL (TANKER)
(RT)
1. 4 VS
VMAT
MSL
Cruising
Ceiling
.5
2.0
DESCENT (D)
1.4 VS
VMAT
MSL
Cruising
Ceiling
.5
2.0
EMERGENCY DESCENT (ED)
1.4 VS
Vmax
MSL
Cruising
Ceiling
.5
2.0
EMERGENCY DECELERATION
(DE)
1.4 VS
Vmax
MSL
Cruising
Ceiling
.5
2.0
AERIAL DELIVERY (AD)
1.2 VS
200 kt
MSL
10,000 ft
0
2.0
TAKEOFF(TO)
Minimum Normal
Takeoff speed
Vmax
MSL
10,000 ft
.5
2.0
CATAPULT TAKEOFF (CT)
Minimum
Catapult End
Airspeed
Vo
MSL
--
.5
nL
APPROACH (PA)
Minimum Normal
Approach Speed
Vmax
MSL
10,000 ft
.5
2.0
WAVE-OFF/GO-AROUND (WO)
Minimum Normal
Approach Speed
Vmax
MSL
10,000 ft
.5
2.0
LANDING (L)
Minimum Normal
Landing Speed
Vmax
MSL
10,000 fl
.5
2.0
* Appropriate to the operational mission.
101
kt
min
+ 30
MIL-STD-1797A
APPENDIX A
If design tradeoffs indicate that significant penalties (in terms of performance, cost, system
complexity, or reliability) are required to provide Level 1 flying qualities in the large envelopes of
Items a-d, above, consideration should be given to restricting the Operational Flight Envelope
toward the minimum consistent with the requirements of the Flight Phase of the operational
mission under consideration.
When effective limiters are employed, Level 1 handling qualities should be provided as close to the limits
as practical. Without an effective limiter, graceful degradation is much preferred in order to preclude a
flying qualities "cliff.”
REQUIREMENT LESSONS LEARNED
Operational missions generally depart significantly from the design mission profile, even for the same
type of mission. It is important to allow enough latitude to cover likely variations. Also, over the life of an
aircraft its operational missions will likely change in both type and detail. There are, of course, tradeoffs
with cost, weight, and the like. For a particular procurement the extent of the Operational Flight Envelope
beyond minimum operational needs should be as large, then, as these trades will reasonably permit.
While stability and control augmentation can do wonders, such factors as basic control authority and rate,
aeroelasticity, and stall speed are (a) limiting at operational extremes and (b) difficult and costly to change
after the design freeze. Skimping on Operational Envelopes, then, can cause difficulties.
There is no connotation that operation is limited to the Operational Flight Envelopes. Operational Flight
Envelopes is a name, the best we could find, for the region in which the best flying qualities are required.
Some pilots have objected that air combat routinely involves flight at lower speeds and higher angles of
attack, even post-stall. It has never been our intent (or indeed within our power) to preclude such
operational use where it is safe. For a particular procurement, requirements outside the Operational Flight
Envelope may warrant strengthening where needs can be identified.
5.1.4 Flight Envelopes-verification
5.1.4.1 Operational Flight Envelopes-verification.
The contractor shall submit the Operational Flight Envelopes for approval by the procuring activity in
accordance with the Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.4.1)
The boundaries of these envelopes will be a competitive consideration, so the contractor can expect to be
held to them once they are established.
VERIFICATION GUIDANCE
Definition of Operational Flight Envelopes is basic to application of the flying qualities requirements.
VERIFICATION LESSONS LEARNED
102
MIL-STD-1797A
APPENDIX A
4.1.4.2 Service Flight Envelopes.
For each Aircraft Normal State the contractor shall establish, subject to the approval of the procuring
activity, Service Flight Envelopes showing combinations of speed, altitude, and normal acceleration
derived from aircraft limits as distinguished from mission requirements. These envelopes shall implicitly
include the ranges of other parameters, such as sideslip, which can be expected within the speed,
altitude and load-factor bounds. For each applicable Flight Phase and Aircraft Normal State, the
boundaries of the Service Flight Envelopes can be coincident with or lie outside the corresponding
Operational boundaries.
REQUIREMENT RATIONALE (4.1.4.2)
The Service Flight Envelope encompasses the Operational Flight Envelope for the same Flight Phase
and Aircraft Normal State. Its larger volume denotes the extent of flight conditions that can be
encountered without fear of exceeding aircraft limitations (safe margins should be determined by
simulation and flight test). At least Level 2 handling qualities are required for normal operation. This
allows a pilot to accomplish the mission Flight Phase associated with the Aircraft Normal State although
mission effectiveness or pilot workload, or both, may suffer somewhat.
This Envelope is also intended to help insure that any deterioration of handling qualities will be gradual as
flight progresses beyond the limits of the Operational Flight Envelope. This serves two purposes. It
provides some degree of mission effectiveness for possible unforeseen alternate uses of the aircraft, and
it also allows for possible inadvertent flight outside the Operational Flight Envelope.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.1.8, 3.1.8.1, 3.1.8.2, 3.1.8.3, and 3.1.8.4.
As with the Operational Flight Envelopes, a range of other flight parameters such as sideslip is implicit.
For the Service Flight Envelope this range should be derived from aircraft limits as distinguished from
mission requirements.
The boundaries of the Service Flight Envelopes are to be based on the speed, altitude, and load factor
considerations discussed below and sketched in figure 7.
1. Maximum Service Speed
The maximum service speed Vmax or Mmax, for each altitude is the lowest of:
a. The maximum speed at which a safe margin exists from any potentially dangerous flight
condition.
b. A speed which is a safe margin below the speed at which intolerable buffet or structural
vibration is encountered.
In setting the maximum service speed, the designer need not consider speed-altitude combinations that
can only be reached in an attitude that would not permit recovery to level flight with a nominal 2000 foot
clearance above sea level while remaining within the Permissible Flight Envelope
2. Minimum Service Speed
The minimum service speed, Vmin or Mmin, for each altitude is the highest of:
a. 1.1 VS
b. VS + 10 knots equivalent airspeed
103
MIL-STD-1797A
APPENDIX A
c.
The speed below which full aircraft-nose-up pitch control power and trim are insufficient to
maintain straight, steady flight.
d. The lowest speed at which level flight can be maintained with MRT.
e. A speed limited by reduced visibility or an extreme pitch attitude that would result in the tail or
aft fuselage contacting the ground.
For engine failure during takeoff, the Standard requires control at speeds down to Vmin (TO); but
requirements for engine-out climb capability are left to performance specifications.
3. Maximum Service Altitude
The maximum service altitude, hmax, for a given speed is the maximum altitude at which a rate of climb of
100 feet per minute can be maintained in unaccelerated flight with maximum augmented thrust (MAT)
4. Service Load Factors
Maximum and minimum service load factors, n(+) [n(-)], are to be established as a function of speed for
several significant altitudes. The maximum [minimum] service load factor, when trimmed for 1 g flight at a
particular speed and altitude, is the lowest [highest] algebraically of:
a. The positive [negative] structural limit load factor.
b. The steady load factor corresponding to the minimum allowable stall warning angle of attack
(4.8.4.2).
c.
The steady load factor at which the pitch control is in the full aircraft-nose-up [nose-down]
position.
d. A safe margin below [above] the load factor at which intolerable buffet or structural vibration
is encountered.
REQUIREMENT LESSONS LEARNED
Rarely has a military aircraft been used only for its design missions. Examples are plentiful: The P-47,
designed as a high-altitude fighter but used more extensively in a ground support role; the B-47 and B-52,
designed for high-altitude penetration and bomb-drop from level flight, but later assigned low-level
penetration and toss delivery as well; the F-4, early assigned a Navy interceptor role but then used as an
all-purpose fighter, and by the Air Force as well; the F-15 and F-16, designed as air superiority fighters
but now also flown heavily loaded with external stores for dive bombing; the C-5, with largely unused
capability for low-level penetration and forward-base operation.
In general, experience strongly indicates the significant benefits to be had from providing at least
acceptable, if not satisfactory, flying qualities up to safe margins from stall, limit dive speed, etc. The
capability to use all the performance fallout outside the design mission envelope can greatly enhance the
worth of any military aircraft. A significant additional benefit is the promotion of graceful degradation
rather than flying qualities “cliffs", for which everything is fine right up to the boundary for departure from
controlled flight.
104
MIL-STD-1797A
APPENDIX A
5.1.4.2 Service Flight Envelopes-verification.
The contractor shall submit the required data for approval by the procuring activity in accordance with the
Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.4.2)
Definition of Service Flight Envelopes is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
The size of Service Flight Envelopes may be a secondary competitive consideration.
VERIFICATION LESSONS LEARNED
105
MIL-STD-1797A
APPENDIX A
4.1.4.3 Permissible Flight Envelopes.
The contractor shall define Permissible Flight Envelopes, subject to the approval of the procuring activity,
which encompass all regions in which operation of the aircraft is both allowable and possible, and which
the aircraft is capable of safely encountering. These Envelopes define boundaries in terms of speed,
altitude, load factor, and any other flight limits.
REQUIREMENT RATIONALE (4.1.4.3)
This is a flight safety consideration.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.9.
Basically, the Permissible Flight Envelope is designed such that, from all points within it, ready and safe
return to the Service Flight Envelope is possible without exceptional pilot skill or technique, regardless of
component or system failures. Structural limits on other flight parameters such as sideslip or lateral
acceleration must also be (a) observed and (b) readily complied with. The requirements on stall, spin, and
dive characteristics, or dive recovery devices, and on approach to dangerous flight conditions also apply,
although Level 2 flying qualities (for Aircraft Normal States) are not required there.
The maximum permissible speed in dives or level flight (the F-103 could reach Vmax in a climb), and the
minimum permissible speed in level flight, can and must be defined for pilots' information. Additionally,
some minimum airspeed may need to be defined for zooms, to assure recoverability. For maneuvers such
as spins, no minimum permissible speed is normally stated; one accepts the low airspeed attained in the
maneuver, satisfactory recovery being the only criterion.
To specify these considerations the contractor must, as a minimum, define the boundaries given below.
1. Maximum Permissible Speed
The maximum permissible speed for each altitude shall be the lowest of:
a. Limit speed based on structural considerations.
b. Limit speed based on engine considerations.
c.
The speed at which intolerable buffet or structural vibration is encountered.
In setting the maximum permissible speed, the designer need not consider speed-altitude combinations
that can only be reached in an attitude that would not permit recovery to level flight with a nominal 2000
foot clearance above sea level while remaining within the Permissible Flight Envelope. To allow for
inadvertent excursions beyond placard speed, some margin should be provided between the maximum
permissible speed and the high-speed boundaries of the Operational and Service Flight Envelopes. Such
a margin is not specified because no satisfactory general requirement could be formulated. However, for
specific designs, the procuring activity should consider 1.1 VH (commonly used for structural
specifications) or the upset requirements of FAR Part 25 and Advisory Circular AC 25-253-IA.
2. 1-g Minimum Permissible Speed
Where maximum lift determines minimum speed, the minimum permissible speed in 1-g flight is VS as
defined in 3.4.2. In general, this is fundamentally more an angle-of-attack limit that an airspeed limit. For
some aircraft, considerations other than maximum lift determine the minimum permissible speed in 1-g
flight [e.g., ability to perform altitude corrections, excessive sinking speed, ability to execute a waveoff
(go-around), etc.]. In such cases an arbitrary angle-of-attack limit, or similar minimum speed and
maximum
106
MIL-STD-1797A
APPENDIX A
load factor limits, shall be established for the Permissible Flight Envelope, subject to the approval of the
procuring activity. This defined minimum permissible speed is to be used as VS in all applicable
requirements. VS needs to be consistent with that used for performance, structure, etc. requirements.
REQUIREMENTS LESSONS LEARNED
For both combat and training missions, flight outside the Service Flight Envelope may well be routine. The
new flying qualities requirements that apply there are largely qualitative. Nevertheless these areas can be
important parts of a useful flight envelope, or may be reached inadvertently, and so need careful
consideration.
5.1.4.3 Permissible Flight Envelopes-verification.
The contractor shall provide the required data for approval by the procuring activity, in accordance with
the Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.4.3)
Definition of Permissible Flight Envelopes is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
The Permissible Flight Envelopes will be refined as better analysis, simulation and test data become
available.
VERIFICATION LESSONS LEARNED
107
MIL-STD-1797A
APPENDIX A
4.1.5 Configurations and States of the aircraft.
The requirements of this standard apply for all configurations required or encountered in the applicable
Flight Phases of 3.2. A selected configuration is defined by the positions and adjustments of the various
selectors and controls available to the crew except for pitch, roll, yaw, throttle and trim controls. Examples
are: the flap control setting and the yaw damper ON or OFF. The selected configurations to be examined
must include those required for performance demonstration and mission accomplishment. Additional
configurations to be investigated are defined as follows: __________. Switches which activate stability
augmentation necessary to meet the requirements of this standard are considered always to be ON
unless otherwise specified. The State of the aircraft is defined by the selected configuration together with
the functional status of each of the aircraft components or systems, throttle setting, weight, moments of
inertia, center-of-gravity position, and external store complement. The trim setting and the position of the
pitch, roll, and yaw controls are not included in the definition of Aircraft State since they are often
specified in the requirements.
REQUIREMENT RATIONALE (4.1.5)
The requirement is intended to assure that all expected aircraft configurations and states are defined and
considered, and that the conditions for compliance are sufficiently called out. All aircraft configurations
either necessary or likely to be encountered must be evaluated.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.1.5 and 3.1.6.
The designer must define the configuration or configurations which his aircraft will have during each Flight
Phase. This includes the settings of such controls as flaps, speed brakes, landing gear, wing sweep, high
lift devices, and wing incidence that are related uniquely to each aircraft design. The requirement
specifies that the configurations to be examined shall include those required for performance
demonstration and mission accomplishment. The position of yaw, roll, pitch, trim controls, and the thrust
setting are not included in the definition of configuration since the positions of these controls are usually
either specified in the individual requirements or determined by the specified flight conditions.
The requirements are stated for Aircraft States and Flight Phases, rather than for aircraft configurations.
The flying qualities should generally be a function of the job to be done rather than of the configuration of
the aircraft. Special considerations or features may require investigation of additional configurations.
These paragraphs introduce the Aircraft State terminology for use in the requirements.
REQUIREMENT LESSONS LEARNED
108
MIL-STD-1797A
APPENDIX A
5.1.5 Configurations and States of the aircraft-verification.
The contractor shall furnish a list of aircraft configurations in accordance with the Contract Data
Requirements List.
VERIFICATION RATIONALE (5.1.5)
Definition of aircraft states is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
This is defined by the requirement itself, and by any specific requirements from the procuring activity.
VERIFICATION LESSONS LEARNED
109
MIL-STD-1797A
APPENDIX A
4.1.6 Aircraft Normal States.
The contractor shall define and tabulate all pertinent items to describe the Aircraft Normal States (no
component or system failure) associated with each of the applicable Flight Phases. This tabulation shall
be in the format and use the nomenclature of table II. Certain items, such as weight, moments of inertia,
center-of-gravity position, wing sweep, or thrust setting may vary continuously over a range of values
during a Flight Phase. The contractor shall replace this continuous variation by a limited number of values
of the parameter in question which will be treated as specific States, and which include the most critical
values and the extremes encountered during the Flight Phase in question.
REQUIREMENT RATIONALE (4.1.6)
Definition of normal aircraft states is basic to application of the flying qualities requirements.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.1.6.1 and 4.2.
It is possible that items not normally considered, such as setting or automatic operation of engine bypass
doors, can affect flying qualities.
The contractor is required to define the Aircraft Normal States for each applicable Flight Phase, in the
format of table II. If the position of any particular design feature can affect flying qualities independently of
the items in table II, its position should be tabulated as well. Initially, variable parameters should be
presented in discrete steps small enough to allow accurate interpolation to find the most critical value, or
combinations for each requirement; then those critical cases should be added. As discussed under 4.1.1
through 4.1.3, center-of-gravity positions that can be attained only with prohibited, failed, or
malfunctioning fuel sequencing need not be considered for Aircraft Normal States.
REQUIREMENT LESSONS LEARNED
5.1.6 Aircraft Normal States-verification.
The contractor shall furnish a list of Aircraft Normal States in accordance with the Contract Data
Requirements List.
VERIFICATION RATIONALE (5.1.6)
Definition of normal aircraft states is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
Definition of normal aircraft states is basic to application of the flying qualities requirements.
VERIFICATION LESSONS LEARNED
110
MIL-STD-1797A
APPENDIX A
4.1.6.1 Allowable Levels for Aircraft Normal States.
Flying qualities for Aircraft Normal States within the Operational Flight Envelope shall be Level 1. Flying
qualities for Aircraft Normal States within the Service Flight Envelope but outside the Operational Flight
Envelope shall be Level 2 or better. To account for the natural degradation of pilot-vehicle performance
and workload in intense atmospheric disturbances, the requirements of 4.1.6.1 through 4.1.6.3 are
adjusted according to 4.9.1.
4.1.6.2 Flight outside the Service Flight Envelopes.
From all points in the Permissible Flight Envelopes and outside the Service Flight Envelopes, it shall be
possible readily and safely to return to the Service Flight Envelope without exceptional pilot skill or
technique. The requirements on flight at high angle of attack, dive characteristics, dive recovery devices
and dangerous flight conditions shall also apply in all pertinent parts of the Permissible Flight Envelopes.
4.1.6.3 Ground operation.
Some requirements pertaining to taxiing, takeoffs, and landing involve operation outside the Operational,
Service, and Permissible Flight Envelopes, as at VS or on the ground. When requirements are stated at
conditions such as these, the Levels shall be applied as if the conditions were in the Operational Flight
Envelopes.
REQUIREMENT RATIONALE (4.1.6.1 - 4.1.6.3)
Levels of flying qualities as indicated in 3.3 apply generally within the Operational and Service Flight
Envelopes. Some basic requirements, generally qualitative in nature, apply in both the Operational and
Service Flight Envelopes. Provision must also be made for expected and allowable operation outside
these envelopes.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.1.10, 3.1.10.1, 3.1.10.3.1, 3.1.10.3.2,
3.1.10.3.3, 3.8.3, and 3.8.3.1.
Aircraft Normal States include both all-up operation and degradations/failures that are sufficiently
probable to be considered Normal. See 4.1.7 and 4.1.7.1 for guidance on the latter.
Note that flying qualities which "warrant improvement" according to figure 6 nevertheless meet all the
requirements if they only occur outside the Operational Flight Envelope.
Where Levels are not specified, care should be taken in selecting requirements from this handbook that
will not overburden the designer. We have tried to keep the impact of 4.1.6.1 in mind in writing the
recommended material to fill in the blanks, but qualitative words such as “objectionable" must be taken in
the context of relevance to operational use.
Since there are few requirements in Aircraft Failure States outside the Service Flight Envelope, implicit
assumptions for 4.1.6.2 are that:
Failures at these conditions are very rare, or
Not-so-rare failures at these conditions are manageable
Given one or more failures within the Service Flight Envelope which would have serious
consequences beyond, at a minimum the crew would be warned away from danger (4.1.8).
Similar assumptions apply for 4.1.6.3. In any given case, their validity will need to be checked.
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MIL-STD-1797A
APPENDIX A
REQUIREMENT LESSONS LEARNED
5.1.6.1 Allowable Levels for Aircraft Normal States-verification.
Verification shall be by analysis, simulation and test.
5.1.6.2 Flight outside the Service Flight Envelopes-verification.
Verification shall be by analysis, simulation, and test.
5.1.6.3 Ground operation-verification.
Verification shall be by analysis, simulation, and test.
VERIFICATION RATIONALE (5.1.6.1 - 5.1.6.3)
These paragraphs are needed to guide the application of the rest of the requirements.
VERIFICATION GUIDANCE
Compliance will be shown in demonstrating compliance with the other requirements of the standard.
VERIFICATION LESSONS LEARNED
112
MIL-STD-1797A
APPENDIX A
4.1.7 Aircraft Failure States.
The contractor shall define and tabulate all Aircraft Failure States which can affect flying qualities. Aircraft
Failure States consist of Aircraft Normal States modified by one or more malfunctions in aircraft
components or systems; for example, a discrepancy between a selected configuration and an actual
configuration. Those malfunctions that result in center-of-gravity positions outside the center-of-gravity
envelope defined in 4.1.1 shall be included. Each mode of failure shall be considered in all subsequent
Flight Phases.
REQUIREMENT RATIONALE (4.1.7)
This tabulation is the starting point for a failure modes and effects analysis, which is necessary in a
complex aircraft to assure flying qualities adequate for mission effectiveness and flight safety.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.6.2.
Because of the exhaustive work often involved and low confidence in the failure probability calculations,
there is a tendency for the procuring activity to substitute a priori list of specific failures. If a design is far
enough along and not excessively complex, such an approach can work. See the guidance for 4.1.7.1.
However, generally comprehensive reliability analyses will be required anyway.
Whether the approach to failure effects on flying qualities is probabilistic, generic or a combination, failure
possibilities of the specific aircraft must be catalogued thoroughly enough to assure adequate mission
effectiveness and flight safety.
REQUIREMENT LESSONS LEARNED
There is more to determining Failure States than just considering each component failure in turn. Two
other types of effects must be considered. First, failure of one component in a certain mode may itself
induce other failures in the system, so failure propagation must be investigated. Second, one event may
cause loss of more than one part of the system or can affect all channels: a broken bracket, a single
crack, a fire, in electrical short, inadequate ground checkout, etc.. The insidious nature of possible
troubles emphasizes the need for caution in design applications.
5.1.7 Aircraft Failure States-verification.
The contractor shall furnish the required data in accordance with the Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.7)
Definition of aircraft failure states is basic to the application of the flying qualities requirements.
VERIFICATION GUIDANCE
Generally, compliance will amount to identifying pertinent items from the list required by the reliability
specification, and checking for completeness. Although the task may seem formidable, the alternative to a
thorough review is a certainty that something important will be overlooked.
VERIFICATION LESSONS LEARNED
113
MIL-STD-1797A
APPENDIX A
4.1.7.1 Allowable Levels for Aircraft Failure States.
Higher performance of aircraft has led to ever-expanding Flight Envelopes, increased control system
complexity, and the necessity to face the problem of equipment failures in a realistic manner. The
specification of Levels corresponding to Failure States is directed at the achievement of adequate flying
qualities without imposing undue requirements that could lead to unwarranted system complexity or
decreased flight safety. For example, an aircraft with two separate pitch controllers is safer from the
standpoint of controller jam but the probability of such a failure is higher. Without actually requiring a
good-handling basic airframe, the standard demands:
High probability of good flying qualities where the aircraft is expected to be used.
Acceptable flying qualities in reasonably likely, yet infrequently expected, conditions.
A floor to assure, to the greatest extent possible, at least a flyable aircraft no matter what failures
occur.
A process to assure that all the ramifications of reliance on powered controls, stability
augmentation, etc., receive proper attention.
Two options are presented to allow the procuring agency to quantitatively specify the allowable
degradation in flying qualities due to failure states. The first option is unchanged from MIL-F-8785C. It
involves the following failure probability calculations:
Identify those Aircraft Failure States which have a significant effect on flying qualities (4.1.7).
Calculate the probability of encountering various Aircraft Failure States, per flight.
Determine the degree of flying qualities degradation associated with each Aircraft Failure State.
Compute the total probability of encountering Level 2 and 3 flying qualities in the Operational
Flight Envelope. This total will be the sum of the probability of each failure if the failures are
statistically independent.
With this method, requirements still remain on the flying qualities effects of certain specific failures, e.g.
engine and flight control system failures.
The second option assumes that certain listed failures and combinations of failures will occur sometime
(with probability 1). As in Option 1, the degraded flying qualities for each selected Failure State are then
evaluated. This approach is referred to as generic failure analysis. Option 2 is provided to allow a formal
Handbook requirement that reflects a current practice. The procuring activity may in fact require
probability calculations for certain axes or system components and a generic failure analysis for others.
The generic failure analysis therefore encompasses the requirements for specific failures of MIL-F-8785C
(3.1.10.2.1)
Generally, the requirements consider only degradations in a single flying quality. It should be recognized
that degradations in several flying qualities parameters can have an effect worse than any one of those
degradations. However, data definitive enough for a standard are not available.
Note that the factors called out in 4.1.11.5 Control margin include Aircraft Failure States and maneuvering
flight appropriate to the Failure States.
114
MIL-STD-1797A
APPENDIX A
4.1.7.2 Aircraft Special Failure States.
Certain components, systems, or combinations thereof may have extremely remote probabilities of failure
during a given flight. The failures may, in turn, be very difficult to predict with any degree of accuracy.
Special Failure States of this type need not be considered in complying with the requirements of this
standard, if justification for considering them as Special Failure States is submitted by the contractor and
approved by the procuring activity.
REQUIREMENT RATIONALE (4.1.7.1 - 4.1.7.2)
Perfection is not a realistic expectation. This requirement is to determine the practical limits in each case.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.6.2.1.
In most cases, a considerable amount of engineering judgment will influence the procuring activity's
decision to allow or disallow a proposed Aircraft Special Failure State. Probabilities that are extremely
remote are exceptionally difficult to predict accurately. Judgments will weigh consequences against
feasibility of improvement or alternatives, and against projected ability to keep high standards throughout
design, qualification, production, use, and maintenance. Meeting other pertinent requirements--MIL-F87242, AFGS-87221, etc.--should be considered, as should experience with similar items. Generally,
Special Failure States should be brought to the attention of those concerned with flight safety.
Note that the required approval of Aircraft Special Failure States, in conjunction with Level 3 floors and
certain requirements that must be met regardless of component or equipment status, can be used as
desired: for example to require a level of stability for the basic airframe, limit use of stick pushers to
alleviate pitch-up, disallow rudder-pedal shakers for stall warning, rule out fly-by-wire control systems,
require an auxiliary power source, force consideration of vulnerability, etc. The procuring activity should
state those considerations they wish to impose, as completely as they can, at the outset; but it is evident
that many decisions must be made subjectively and many will be influenced by the specific design.
Several categories of Special Failure States can be distinguished. Certain items might be approved more
or less categorically:
•
Control-stick fracture.
•
Basic airframe or control-surface structural failure.
•
Dual mechanical failures in general.
Regardless of the degree of redundancy, there remains a finite probability that all redundant paths will
fail. A point of diminishing returns will be reached, beyond which the gains of additional channels are not
worth the associated penalties:
•
Complete failure of hydraulic or electrical, etc., systems.
•
Complete or critical partial failure of stability augmentation that has been accepted as
necessary to meet Level 3.
Some items might be excepted, if special requirements are met. For example, some limited control should
remain after failure of all engines, provided by accumulators or an auxiliary power source as appropriate.
In the last analysis the procuring activity is responsible for approving design tradeoffs that bear upon
safety. Rather than inhibiting imaginative design, then, this paragraph should be construed as forcing
examination of failure possibilities as they affect flight safety through deterioration of flying qualities. The
present state of the art can support some properly implemented reliance on stability augmentation to
maintain Level 3 flying qualities, but it must be done carefully and for good reason.
115
MIL-STD-1797A
APPENDIX A
Concerning the admissibility of a Special Failure State on the basis of its remoteness of possibility, the
combined probability of having any flying qualities worse than Level 3--not just each individual Failure
State probability--must be kept extremely remote.
REQUIREMENT LESSONS LEARNED
By default, all failure modes not considered become Special Failure States, albeit without specific
approval
5.1.7.1 Allowable Levels for Aircraft Failure States-verification
5.1.7.2 Aircraft Special Failure States-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirements List, for
review by the procuring activity.
VERIFICATION RATIONALE (5.1.7.1 - 5.1.7.2)
Definition of Aircraft Special Failure States is basic to application of the flying qualities requirements.
VERIFICATION GUIDANCE
Definition of Aircraft Special Failure States is basic to application of the flying qualities requirements.
VERIFICATION LESSONS LEARNED
116
MIL-STD-1797A
APPENDIX A
4.1.7.3 Probability calculation.
When Aircraft Failure States (4.1.7) exist, a degradation in flying qualities is permitted only if the
probability of encountering a lower Level than specified is sufficiently small. The contractor shall
determine, based on the most accurate available data, the probability of occurrence of each Aircraft
Failure State per flight within the Operational and Service Flight Envelopes. These determinations shall
be based on ____________except that:
a. All aircraft systems are assumed to be operating for the entire flight, unless clearly operative
only for a shorter period
b. For these calculations, the length of flight shall be hours
c.
Each specific failure is assumed to be present at whichever point in the Flight Envelope being
considered is most critical (in the flying qualities sense).
From these Failure State probabilities and effects, the contractor shall determine the overall probability,
per flight, that one or more flying qualities are degraded to Level 2 because of one or more failures. The
contractor shall also determine the probability that one or more flying qualities are degraded to Level 3.
These probabilities shall be less than the values shown in table III.
TABLE III. Levels for Aircraft Failure States.
PROBABILITY OF
ENCOUNTERING
WITHIN OPERATIONAL
FLIGHT ENVELOPE
Level 2 after failure
< ______ per flight
Level 3 after failure
< ______ per flight
WITHIN SERVICE FLIGHT
ENVELOPE
< ______ per flight
REQUIREMENT RATIONALE (4.1.7.3)
This requirement provides a sound analytical method for accounting for the effects of failures. It serves to
force a detailed failure mode and effect analysis from the flying qualities standpoint. Such an analysis is
vital as both system complexity and the number of design options increase.
The probability of a degraded Level of flying qualities is related to, but not exactly the same as, mission or
flight-safety reliability. A degraded flying qualities Level is allowed for some infrequently expected events
failure of aircraft systems or flight outside the Operational Flight Envelope, near the aircraft's limits (by
definition, the Operational Flight Envelopes encompass the design missions).
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.1.10.2, which calls for determinations to be based
on MIL-STD-756.
As in MIL-F-8785C, failure probabilities are specified as a function of number of flights, rather than flight
hours. As discussed in "Supporting Data," a typical flight time of four hours was used for the MIL-F-8785B
numbers. The numbers in table VIII, while arbitrary, were chosen so that failures which degrade flying
qualities would not contribute disproportionately to reduction or loss of mission effectiveness, or to flight
safety problems. The form is consistent with failure rate data, which usually are presented per flight hour,
117
MIL-STD-1797A
APPENDIX A
and with the critical takeoff and landing Flight Phases, which occur once per flight. It is implied that, while
mission duration varies among aircraft types, the number of missions does not. The flight length of a
normal long mission, with in-flight refueling if appropriate, should be specified.
TABLE VIII. Recommended Levels for Aircraft Failure States.
PROBABILITY OF
ENCOUNTERING
WITHIN OPERATIONAL
FLIGHT ENVELOPE
WITHIN SERVICE FLIGHT
ENVELOPE
Level 2 after failure
<10-2 per flight
-------
Level 3 after failure
<10-4 per flight
<10-2 per flight
For comparison, MIL-F-9490 recommends the following unreliability allowance for the entire flight control
system, manual and automatic, with somewhat different ground rules:
Mission accomplishment:
< 10-3 per flight
Flight safety:
< 5 x 10-7 per flight, Class III;
< 10-5 per flight, Classes I, II, IV
FAR Part 25 paragraph 25.671, by comparison for the flight control system:
Probable malfunctions > 10-3 per hour are allowed to have only minor effect
Extremely improbable failures < 10-9 per hour need not be considered
Continued safe flight and landing must be assured after all other failures/combinations.
Limited degradation of flying qualities (e.g., Level 1 to Level 2) is acceptable if the combined probability,
of such degradation is small. If the probability of any particular failure is high, then that failure must
produce no degradation beyond the Level required for Normal States. Another way of stating this would
be that in the Operational Envelope the probability of encountering Level 2 any time at all on a given flight
should not, according to the table VIII recommendations, exceed 10-2 and the probability of encountering
Level 3 on any portion of the flight should not exceed 10-4. Somewhat reduced requirements are to be
imposed for flight within the Service Flight Envelope, for both Normal and Failure States. Outside the
Service Flight Envelope, most of the requirements of the standard do not apply. There is, however, a
qualitative requirement to be able to return to the Service Flight Envelope after a failure (i.e., Paragraph
4.1.7.6).
The numbers are given as orders of magnitude. When predicting the occurrence of events of such small
probability, that is about the most accuracy that can be expected.
The requirements do not account for the expectation that degradation of more than one flying quality will
be more severe than any of those degradations singly. In the absence of a definitive data base,
simulation is recommended.
Degradation in atmospheric disturbances is discussed in Requirement Guidance for 4.9.1
The probability of flying qualities degradation is influenced by a number of factors such as design
implementation and complexity (including reconfiguration capability), computer reliability improvements,
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MIL-STD-1797A
APPENDIX A
lightning protection, built-in test (BIT), maintenance practices and dispatch rules. Peacetime vs. wartime
operation can be a necessary concern, although battle damage is a separate consideration.
The numerical values in table III should reflect specific requirements for a given weapon system. The
procuring-activity engineer should, as a matter of course, confer with both the using-command
representative and the reliability engineers to assure that the probabilities associated with the Levels are
consistent with the overall design goals. However, the recommended values of table IV are reasonable,
based on experience with past aircraft and current and projected states of the art. To illustrate this, the
following listing presents actual control system failure information for several piloted aircraft:
Reference
System
Mean Time Between
Malfunctions (MTBM)
Bureau of Naval Weapons
Failure Rate Data Handbook
F-101B
F-104
F-105D(Fit cntl + elect)
E-1B
86 hours
300 hours
14 hours
185 hours
B-58
20 hours
F-16A
F-16B
F-16C
F-16D
KC-10A
A-10A
F-4C
F-4D
F-4E
F-4G
F-15C
B-1
48 hours
40 hours
66 hours
68 hours
130 hours
70 hours
26 hours
22 hours
22 hours
17 hours
49 hours
8 hours
Ad Hoc Committee Report on
B-58 Controllability in Flight
MODAS
MODAS stands for Maintenance and Operational Data Access Systems, a system the Air Logistics
Centers use to record and document failures of aircraft systems. The tabulated MODAS data from early
1986 are for type 1 flight control system failures only, not including the autopilot.
Unfortunately the flying qualities effects of the reported failures are not given along with the above data. A
Second Analysis of B-58 Flight Control System Reliability indicates, however, that the mean time between
critical failures is about five times the MTBM. If critical failures are ones that degrade one or more flying
qualities to Level 2, then for a typical average flight time of four hours:
119
MIL-STD-1797A
APPENDIX A
P(Level 2)
This yields:
=
Probability of encountering Level 2 flying qualities during a single flight
=
1- e-4/[5(MTBM)]
=&
4
5(MTBM)
System
F-101B
F-104
F-105D
E-1B
B-58
F-16A
F-16B
F-16C
F-16D
KC-10A
A-10A
F-4C
F-4D
F-4E
F-4G
F-15C
B-1
P(Level 2)
0.0093
0.0027
0.0570
0.0043
0.0400
0.0167
0.0200
0.0121
0.0118
0.0062
0.0114
0.0308
0.0364
0.0364
0.0471
0.0163
0.1000
These data indicate that most systems have P(Level 2) ≈ 10-2 or less (or approximately one out of a
hundred flights). We consider the F-16, A-10, and F-15 to meet the requirement, considering the limited
accuracy of probability calculations. For the F-105, F-4, and B-1, the data may include failures in
electronic components which do not result in degradation of flying qualities. Numbers of roughly the same
magnitude have been used for both American (Tentative Airworthiness Objectives and Standards for
Supersonic Transport Design Proposals) and Anglo-French (Supersonic Transport Aeroplane Flying
Qualities) supersonic transport design.
A more significant analysis was conducted on the F-4 by the Air Force Flight Dynamics Laboratory
(AFFDL-FGC-TM-71-7). The level of degradation used in this report was based on whether or not the
failure resulted in an abort. Failures without abort were considered degraded to Level 2, and those which
caused an abort were considered degraded to Level 3. The results showed that the F-4 handling
qualities, in an average 2.57 hour flight, will be degraded to Level 2 on an average of 0.043/flight, and to
Level 3 a maximum of 0.0021/flight (21 x 10-4).
A similar comparison can be made between accident loss rates and the requirement for P(Level 3) < 2.5
10-5/flight hour. The Level 3 boundaries are, while not necessarily totally safe, at least safety related.
General Dynamics Rpt FZM-12-2652 indicates the following aircraft accident loss rates during 1967. Also,
shown is the probability of aircraft loss, per 4-hour flight, for an assumed exponential loss distribution.
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APPENDIX A
Aircraft
1967 loss rate
(Losses/100,000 hr)
probability of loss
during 4-hour flight
F-101
15
6 x 10-4
F-104
23
9.2 x 10-4
F-105
17
6.8 x 10-4
F-106
10
4 x 10-4
14.1
5.64 x 10-4
F-102
9
3.6 x 10-4
F-100
10
4 x 10-4
Avg
14
5.6 X 10-4
F-4
If Level 3 represented a safety problem, which it conservatively does not, then the allowable 10-4
probability of encounter per 4-hour flight would account for about 1/4 to 1/9 of the total probability, of
aircraft loss. That is, flying-qualities-oriented losses would represent about 1/4 to 1/9 of all losses. Other
losses could be due to engine failures, etc. Therefore, based on experience the recommended table VIII
value is reasonable.
As a final note, "Industry Observer" from Aviation Week and Space Technology of 1 April 1968 indicates
an Army aircraft accident rate of 22.2/100,000 hours which is very close to the previously cited
experience with a number of Air Force aircraft.
REQUIREMENT LESSONS LEARNED
The following excerpts were taken from written comments made by ASD regarding lessons learned
utilizing Paragraph 3.1.10.2 of MIL-F-8785B.
F-16:
Levels were applied to failures without calculating probabilities;
assumed that if a failure could occur, it would eventually (i.e., generic
failure analyses)
F-15, F-16, F-105:
Low confidence in failure probability calculations; better to consider
individual failures (i.e., generic failure analyses)
B-1/AMST:
See ASD-TR-78-13 for approach to failure states taken on B-1 and
AMST. Probability analysis was used. B-1 experience with its longest
mission (10-hours) indicated that meeting the probability of
encountering Level 3 of 10-4/flight [as required in MIL-F-8785B] was not
possible at that time, with that design concept.
F-15, A-10:
Very hard to determine realistic probabilities; recommend defining
special failure states from past experience (i.e., generic failure
analyses)
A-10:
Only specific failures were investigated (i.e., generic failure analyses);
most of front section of specification not really used.
F-5E:
Flight outside the Operational Flight Envelope should not be considered
abnormal; Paragraph 3.1.10.2 of MIL-F-8785B deleted as useless.
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APPENDIX A
See also REQUIREMENT LESSONS LEARNED for 4.1.7.4.
Other problems with this approach are:
a. The scatter in the MTBM data for various components can provide large uncertainty in the
calculation of probability of failure in the system.
b. Systems change all through the design, ground and flight test process. Therefore, the number of
components and their arrangement change significantly. This approach requires that failure possibilities
and modes, and probabilities of failure, be reevaluated after each design change.
5.1.7.3 Probability calculation-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.7.3)
This approach addresses the reliability of flying qualities, rather than the reliability of hardware and
software per se.
Until some time downstream in the design process, the flight control and related subsystems will not be
defined in enough detail to permit a comprehensive listing of failure possibilities, much less to estimate
their likelihood. Initially, this requirement serves as guidance in selecting a design approach and
components and redundancy levels which can potentially achieve or surpass the stated probabilities of
not encountering the degraded Levels. As the design progresses, reliability analyses and failure mode
and effect analyses will provide the means of determining compliance.
VERIFICATION GUIDANCE
The discussions that follow are taken from AFFDL-TR-69-72.
Implementation of the Level concept involves both reliability analyses (to predict failure probabilities) and
failure effect analyses (to insure compliance with requirements). Both types of analyses are in direct
accord with, and in the spirit of, MIL-STD-756B and MIL-STD-882B. These related specifications are, in
turn, mandatory for use by all Departments and Agencies of the Department of Defense. Implementation
of the flying qualities specification is, for the most part, a union of the work required by these related
specifications with normal stability and control analysis.
Failure States influence the aircraft configurations, and even the mission Flight Phases, to be considered.
All failures that could have occurred previously must be examined, as well as all failures which might
occur during the Flight Phase being analyzed. For example, failure of the wings to sweep forward during
descent would require consideration of a wings-aft landing that otherwise would never be encountered.
There are failures that would always result in an aborted mission, even in a war emergency. The pertinent
Flight Phases after such failures would be those required to complete the aborted (rather than the
planned) mission. For example, failure of the flaps to retract after takeoff might mean a landing with flaps
at the takeoff setting, with certain unexpended external stores; but supersonic cruise would be
impossible. If the mission might be either continued or aborted, both contingencies need to be
considered.
There are some special requirements pertaining to failure of the engines and the flight control system. For
these special requirements the pertinent failure is assumed to occur (with a probability of 1), with other
failures considered at their own probabilities. For all other requirements, the actual probabilities of engine
and flight control system failure are to be accounted for in the same manner as for other failures.
Note that specific Special Failure States (4.1.7.2) may be approved; these Failure States need not be
considered in determining the probability of encountering degradation to Level 3. This allows each
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MIL-STD-1797A
APPENDIX A
catastrophic failure possibility to be considered on its own. Requiring approval for each Special Failure
State gives the procuring activity an opportunity to examine all the pertinent survivability and vulnerability
aspects of each design. Survivability and vulnerability are important considerations, but it has not yet
been possible to relate any specific flying qualities requirements to them.
For electronic components, MIL-HDBK-217 provides reliability data. There seems to be no standard
source for reliability data on other components.
A typical approach (but not the only one) for the system contractor is outlined below. The stages indicated
are appropriate for the required calculation and submittal.
1. Initial Design
The basic airframe is designed for a Level 1 target in respect to most flying qualities in the Operational
Flight Envelope. It may quickly become apparent that some design penalties would be inordinate
(perhaps to provide sufficient aerodynamic damping of the short period and dutch roll modes at high
altitude); in those cases the basic airframe target would be shifted to Level 2. In other cases it may be
relatively painless to extend some Level 1 flying qualities over the wider range of the Service Flight
Envelope. Generally the design will result in Level 1 flying qualities in some regions and, perhaps, Level 2
or Level 3 in others. Augmentation of one form or another (aerodynamic configuration changes, response
feedback, control feed forward, signal shaping, etc.) would be incorporated to bring flying qualities up to
Level 1 in the Operational Flight Envelope and to Level 2 in the Service Flight Envelope.
2. Initial Evaluation
The reliability and failure mode analyses are next performed to evaluate the nominal system design
evolved above. All aircraft subsystem failures that affect flying qualities are considered. Failure rate data
for these analyses may be those specified in the related specifications, other data with supporting
substantiation and approval as necessary, or specific values provided by the procuring agency. Prediction
methods used will be in accordance with related specifications. The results of this evaluation will provide:
a) a detailed outline of design points that are critical from a flying qualities/flight safety standpoint; b)
quantitative predictions of the probability of encountering Level 2 in a single flight within the Operational
Flight Envelope, Level 3 in the Operational Envelope, and Level 3 in the Service Envelope; and c)
recommended airframe/equipment changes to improve flying qualities or increase subsystem reliability to
meet the specification requirements. It should be noted that the flying qualities/flight safety requirements
are concerned with failure mode effects, while MIL-STD-785 provides “basic" reliability requirements per
se (all failures regardless of failure effects). In the event of a conflict, the most stringent requirement
would apply.
3. Re-Evaluation
As the system design progresses, the initial evaluation is revised at intervals. This process continues
throughout the design phase, with review by the procuring activity at times consonant with other reviewer
activity.
The results of the analyses of vehicle flying qualities/flight safety may be used directly to: a) establish
flight test points that are critical and should be emphasized in the flight test program; b) establish pilot
training, requirements for the most probable, and critical, flight conditions; and c) provide guidance and
requirements for other subsystem designs.
The failure modes and effects analysis (MIL-STD-1629) will highlight items which need to be checked by
piloted simulation and flight test--although safety considerations may limit flight test. Further, compliance
is demonstrated on the basis of the probability calculations and checked as accumulated flight experience
permits. All of the assumptions, such as independence of failure modes, etc., should be firmly established
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MIL-STD-1797A
APPENDIX A
and mutually agreed upon by the contractor and SPO. The combined effects of turbulence and failures
should also be considered. The boundaries given in figure 8 have been suggested as guidelines for these
combined effects, but the recommended approach is to relax the requirements according to 4.9.1.
The combined effects of failures and turbulence should be validated in a manned simulation. Multiple-axis
failures should also be simulated, especially where the flying quality parameters result in pilot ratings near
the applicable Level's lower limits.
Proof of compliance is, for the most part, analytical in nature as far as probabilities of failure are
concerned. However, some failure rate data on the actual flight equipment may become available during
final design phases and during flight test, and any data from these or other test programs should be used
to further demonstrate compliance. Stability and control data of the usual type (e.g., predictions, wind
tunnel, flight test) will also be used to demonstrate compliance. Finally, the results of all analyses and
tests will be subject to normal procedures of procuring agency approval.
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MIL-STD-1797A
APPENDIX A
FIGURE 8. Definition of Levels which include atmospheric disturbances
as well as failures -- suggested by Carlson (AFFDL-TR-78-171).
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.7.4 Generic failure analysis.
The requirements on the effects of specific types of failures. for example propulsion or flight control
system, shall be met on the basis that the specific type of failure has occurred, regardless of its
probability of occurrence. The requirements of this standard on failure transients shall also be met. The
allowable flying quality Levels for each of the Failure States in 4.1.7 are defined as follows:
_____________. In addition, flying qualities in the following specific failure states shall be as follows:
Failure
Level
_______
______
REQUIREMENT RATIONALE (4.1.7.4)
In accordance with the lessons learned of 4.1.7.3, this paragraph has been included to provide a way to
specify the allowable degradation in handling qualities due to failures without making detailed probability
calculations.
This approach assumes that a given component, or series of components, will fail. Furthermore. the
failures are assumed to occur in the most critical flight condition; for example, a yaw damper failure at the
maximum service ceiling in turbulence. Based on the comments made by users of MIL-F-8785B (see
Lessons Learned in 4.1.7.3), this approach is a common current practice.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.1.10.2.1.
The selection of failures to be considered is based on preliminary estimates of handling quality
degradations. For example, the loss of one to three channels of a quad-redundant SCAS may have no
effect. Conversely the failure of a single-channel, limited-authority damper would warrant a complete
analysis, possibly simulation too, to determine the resulting degradation in flying qualities. Requirements
such as two-fail-operate assume a certain degree of reliability and so may penalize either the
manufacturer or the user. In addition, the procuring activity may wish to demand that certain specific
failures be considered regardless of their probabilities.
Because the selection of failure modes is highly dependent on the details of the design, close
coordination between the contractor and the procuring activity will be required when identifying failure
modes to be analyzed and determining the allowable degradation. Indeed, this is currently practiced.
A comprehensive failure mode and effect analysis has been found essential for all but very simple
designs. The questions here are how to force this analysis and how to treat the flying qualities aspects.
When writing a generic failure requirement, it is best to associate the required Levels of flying qualities
with the number of failures in the system, and the tasks that must be performed, e.g.:
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MIL-STD-1797A
APPENDIX A
1.
2.
3.
4.
5.
6.
7.
Failure State
After two independent failures in the
stability/control augmentation system
After two independent failures in the air data
system
After two independent failures in the electrical
power system
After loss of all propulsion power, means shall be
provided to maintain stable and controlled flight for
the time required to descend from cruise altitude to
sea level at the speed for best L/D, with a 5 minute
reserve
After 2 independent failures in a system (such as a
fuel system) which can affect center of gravity
position
Electrical power interrupts or transients shall not
result in excessive aircraft transients or result in
loss of controlled flight
After 1 failure in the hydraulic system
Flight Phase
All
Level
1
RR,CR,PA,L
2
RR,CR,PA,L
2
ED,L
2
All
2
All
2
All
2
It must be emphasized that this is only an example. There may be failures that are not discussed in this
example, such as generic software faults; Levels of flying qualities for these failure modes should be
negotiated with the contractor and coordinated with each of the technical disciplines involved.
The reader is directed to the specific flying qualities requirements which must be met in the event of
failures. Paragraphs 4.2.6.1, 4.2.8.6.5, 4.5.7.1, 4.5.9.5.6, and 4.6.7.9 concern flight control system
failures; while 4.5.8.4, 4.5.9.5.5, 4.6.5.1, 4.6.6.2 and 4.6.7.8 concern engine failures. Requirements on
failure transients include 4.1.12.8, 4.2.6.1, 4.5.7.1 and 4.6.5.2.
REQUIREMENT LESSONS LEARNED
This approach has been utilized on the F-15 (a Level 2 basic airframe was specified), F-16 (a minor flight
control system change from the prototype), A-10 (a simple flight control system design), and F-5E (a
design evolution) with reasonable success.
As an example of this process, in one case a failure requirement for the hydraulic system was stated as:
"After loss of one hydraulic system, flying qualities sufficient to return to base and land shall remain (Level
2)". A flying qualities analysis was performed to determine the control capability with a reduced number of
control surfaces. It was found that if one hydraulic system provided power to an aileron on one side of the
vehicle, the elevator on the opposing side, and the rudder, sufficient control could be maintained to land
the vehicle. This was confirmed on a piloted simulator. The hydraulic system was plumbed accordingly
and demonstrated on the Iron Bird. As the flight test program progressed, modifications were made to the
flight control laws and the aerodynamic data package. Furthermore, there were also some slight changes
made to the actuators. At the end of the test period, the flying qualities were reevaluated on the simulator
to ensure that the vehicle could be landed. This data was then used to verify specification compliance.
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MIL-STD-1797A
APPENDIX A
5.1.7.4 Generic failure analysis-verification.
The contractor shall submit the required data in accordance with the Contract Data Requirements List.
VERIFICATION RATIONALE (5.1.7-4)
Verification of generic failure analysis is required to demonstrate that flying qualities parameters in
question fall within prescribed boundaries for specified levels when generic failures, as specified in
4.1.7.4, occur.
VERIFICATION GUIDANCE
In most cases, demonstration of compliance will consist of showing that the flying qualities parameters in
question fall within the prescribed boundaries for specified Levels. Where numerical boundaries are not
available, either ground-based or- in-flight simulation will be required. Failures in more than one axis that
cause the specified flying quality parameters to fall near the boundaries of more than one flying qualities
requirement should also be simulated. Finally, the combined effects of failures and turbulence should be
investigated utilizing a piloted simulation.
VERIFICATION LESSONS LEARNED
See 4.1.7.3.
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MIL-STD-1797A
APPENDIX A
4.1.7.5 When Levels are not specified.
Within the Operational and Service Flight Envelopes, all requirements that are not identified with specific
Levels shall be met under all conditions of component and system failure except approved Aircraft
Special Failure States (4.1.7.2).
REQUIREMENT RATIONALE (4.1.7.5)
A number of basic requirements, generally qualitative in nature, apply regardless of Failure State, in both
the Operational and Service Flight Envelopes.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.10.3.2.
In view of this necessary gap-filling requirement, care should be taken in selecting requirements from this
Handbook that will not overburden the designer. We have tried to keep the impact of 4.1.7.5 in mind in
writing the recommended material to fill in the blanks, but qualitative words such as “objectionable" must
be taken in the context of the relative likelihood of failure and relevance to operational use.
REQUIREMENT LESSONS LEARNED
5.1.7.5 When Levels are not specified-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.1.7.5)
A failure modes and effects analysis will show conditions critical to particular requirements.
VERIFICATION GUIDANCE
Verification will require qualitative evaluation.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.7.6 Failures outside the Service Flight Envelopes.
Failures to be considered outside the Service Flight Envelopes but within the corresponding Permissible
Flight Envelopes are __________. After these failures it shall be possible to return safely to the Service
and Operational Flight Envelopes.
REQUIREMENT RATIONALE (4.1.7-6)
Air combat may require short excursions to angles of attack beyond stall, or the normal limit angle of
attack, frequently enough to warrant special consideration. Other tasks and ground operation may be
pertinent.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.10.3.3.
The other post-stall requirements rely on the basic airframe, the flight control system specification, the
presumed small chance of failure in the post-stall region, and avoiding the region if certain Aircraft Failure
States exist. New emphasis on post-stall technology now gives added importance to this flight region for
certain combat aircraft. At this stage the only guidance possible is to raise the issue and to list some
factors in avoiding loss of control:
Engine: flameout, duty cycle or throttle usage, compressor stall
Reaction controls
Fail operate or fail soft
Frequency of failure
Failure-warning reliability.
REQUIREMENT LESSONS LEARNED
As for normal operation, stability-augmentation failure modes can have completely different effects in the
post-stall region from their action in normal flight envelopes.
5.1.7.6 Failures outside the Service Flight Envelope-verification.
The contractor shall review the list furnished by the procuring activity and through its own analysis modify
and extend that list as necessary for adequate coverage of flying qualities degradations, subject to
procuring activity approval, in accordance with the Contract Data Requirements List. Verification of safe
return capability shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.1.7-6)
Flight safety will be a prime factor in determining the means of verification.
VERIFICATION GUIDANCE
Outside the Service Flight Envelope, wind tunnel tests, structural analyses, propulsion limits, etc., will
need to be taken into account.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.8 Dangerous flight conditions.
Dangerous conditions may exist at which the aircraft should not be flown. When approaching these flight
conditions, it shall be possible by clearly discernible means for the pilot to recognize the impending
dangers and take preventive action. Whenever failures occur that require or limit any flight crew action or
decision concerning flying the aircraft, the crew member concerned shall be given immediate and easily
interpreted indication.
4.1.8.1 Warning and indication.
Warning and indication of approach to a dangerous condition shall be clear and unambiguous. For
example, a pilot must be able to distinguish readily among stall warning (which requires pitching down or
increasing speed), Mach buffet (which may indicate a need to decrease speed), and normal aircraft
vibration (which indicates no need for pilot action).
4.1.8.2 Devices for indication, warning, prevention, and recovery.
It is intended that dangerous flight conditions be eliminated and the requirements of this standard be met
by appropriate aerodynamic design and mass distribution, rather than through incorporation of a special
device or devices. As a minimum, these devices shall perform their function whenever needed but shall
not limit flight within the Operational Flight Envelope. Neither normal nor inadvertent operation of such
devices shall create a hazard to the aircraft. For Levels 1 and 2, nuisance operation shall not be possible.
Functional failure of the devices shall be indicated to the pilot.
REQUIREMENT RATIONALE (4.1.8 - 4.1.8.2)
Approach to any dangerous flight condition must be clearly apparent to the pilot with sufficient margin
(time, control power, etc.) to avoid loss of control. That, together with limiting the frequency of encounter,
is the essence of flight safety as it involves flying qualities.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.4.1, 3.4.1.1, 3.4.1.2 and 3.4.9.
The need for warning may not become apparent until late in the development program (or after it), and
each such device will generally have to be tailored to a specific set of conditions. These requirements
clearly apply to stall warning and prevention devices, as well as to other types.
Certain failures may require restriction of the flight envelope in order to assure safety after the failure, or
in the event of a subsequent failure. The flight crew needs to be made aware when such a situation
exists.
One possible source of danger is the use of special, unconventional control modes for certain tasks.
Under stress, a pilot may forget and revert to the normal technique or exceed a limit. For example, a yaw
pointing mode might generate excessive sideslip at low dynamic pressure or too much lateral
acceleration at high dynamic pressure. Also, partially or fully automatic modes of flight need to be
examined for possible hazards.
Normally, a reasonably reliable limiter or warning need not be redundant. If the pilot knows that the device
is inoperative, he can stay well clear of the danger. Nuisance operation not only interferes with mission
performance; it breeds disregard or disconnection. Reliance is placed on the flight control system
specification and other specifications and standards to assure sufficient reliability of these devices,
warning of their failure, and checkout provisions.
Requirement 4.1.8.2 is designed to discourage prevention devices that create more problems than they
solve.
Testing will be necessary to assure that the limiting or warning is satisfactory in all maneuvers and that
functional failure of any such devices is indicated to the pilot. Ultimately, flight testing will be required
(see, for example, MIL-S-83691).
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MIL-STD-1797A
APPENDIX A
REQUIREMENT LESSONS LEARNED
Stall limiters have proved to be of significant help, as with the F-101B Boundary Control System and the
F-16 stall, load factor and roll limiter, which allow carefree maneuvering up to the set limits. However, an
undefeatable limiter is hard to design. If a dangerous pitch-up or locked-in deep stall lurks beyond, some
pilot will encounter it if that's possible. Indeed pilots would rather bend the wings than hit the ground, so a
soft limiter may be in order. On the other hand, makers and flyers of aircraft with no post-stall limitations
(e.g., T-38/F-5, F-15) find these extreme angles of attack useful occasionally in air combat.
Several C-133 losses over oceans were conjectured to have resulted from starting long-range cruise too
close to stall, with no stall warning and a severe rolloff in a power-on stall. The artificial stall warning often
was turned off because it was not reliable.
F-16s have been lost in ground attack runs because of deteriorated aerodynamics with external stores.
The fix was to change the limits when air-to-ground stores are carried. Although the F-16 uses a pilotoperated switch, this can be done automatically, so that the pilot does not have to remember to switch.
5.1.8 Dangerous flight conditions-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
5.1.8.1 Warning and indication-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
5.1.8.2 Devices for indication, warning, prevention, and recovery-verification.
Verification shall be by analysis, simulation, and ground and flight testing.
VERIFICATION RATIONALE (5.1.8 - 5.1.8.2)
In the end, the procuring and test activities must assess the degree of danger and their test pilots must
pass on the acceptability of devices and aircraft characteristics.
VERIFICATION GUIDANCE
FTC-TIH-79-2, FTC-TD-73-2, and USNTPS-FTM-103 contain guidance on flight testing for stall/post-stall
and other conditions of concern. MIL-S-83691 is an Air Force flight test specification and MIL-D-8708 the
corresponding Navy specification.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.9 Interpretation of subjective requirements.
In several instances throughout the standard, subjective terms, such as objectionable flight
characteristics, realistic time delay, normal pilot technique and excessive loss of altitude or buildup of
speed, have been employed where insufficient information exists to establish absolute quantitative
criteria. Final determination of compliance with requirements so worded will be made by the procuring
activity.
REQUIREMENT RATIONALE (4.1.9)
This statement is included to clarify the authority of the procuring activity in determining compliance with
subjective requirements. These requirements bear on significant mission effectiveness and flight safety
considerations in areas where data are lacking to establish quantitative requirements.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.11.
While subjective requirements permit wide latitude for the contractor in the early stages, the focus in the
flying qualities specifications has been, and will continue to be, on quantifying all requirements for which
sufficient data exists. The procuring activity should always have final power in accepting compliance with
subjective requirements.
REQUIREMENT LESSONS LEARNED
Validity of ground-based and in-flight simulations should be examined as closely as possible. For
example it is not entirely clear whether failure of moving-base simulation to predict YF-17 pilot-induced
oscillation tendencies prior to flight was due to simulator limitations or inaccurate representation of the
flight control system as actually flown. In-flight simulation inaccuracies can also cause discrepancies, viz.
the NT-33 simulation which predicted worse AFTI-F-16 ratings than the pilots actually gave the airplane.
The cause was traced to modeling error. The dynamics of ground-based simulators (motion, visual scene,
computation time, etc.) in general are poorly documented.
5.1.9 Interpretation of subjective requirements-verification.
Verification shall be by analysis, simulation and test.
VERIFICATION RATIONALE (5.1.9)
Verification will require qualitative evaluation.
VERIFICATION GUIDANCE
The procuring activity will rely heavily upon comments by evaluation pilots during simulations, and later
during flight tests, in determining compliance with subjective requirements.
In a draft of flying qualities for military rotorcraft, Hoh, and others, supplement the requirements on
rotorcraft characteristics with required pilot ratings plus desired and adequate performance in a list of task
related demonstration maneuvers (rapid hovering turn, lateral unmask and remask, etc.). These
maneuvers, which are spelled out in some detail, are to be performed by several pilots. While such added
requirements would be of limited help in the design stage, they would verify the suitability of both
qualitative and quantitative flying qualities for the intended missions.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.10 Interpretation of quantitative requirements.
Many of the numerical requirements of this standard are stated in terms of a linear mathematical
description of classical aircraft. Certain factors, for example flight control system nonlinearities and higherorder dynamics or aerodynamic nonlinearities, can cause an appreciable difference in the aircraft
response apparent to the pilot from that of the linear model of the basic airframe. The contractor shall
determine equivalent classical systems which have responses most closely), matching those of the actual
aircraft. Then those numerical requirements of section 4 which are stated in terms of linear system
parameters (such as frequency, damping ratio and modal phase angles) apply to the parameters of that
equivalent system rather than to any particular modes of the higher-order system representation. The
adequacy of the response match between equivalent and actual aircraft, or alternative criteria, shall be
agreed upon by the contractor and the procuring activity. Nonlinearities or higher-order dynamics that
may exist shall not result in any objectionable (for Levels 1 and 2) or dangerous characteristics.
REQUIREMENT RATIONALE (4.1.10)
Where higher-order effects are involved, consideration of any particular roots can be misleading. In
acknowledgment of the increasing complexity of aircraft control systems, equivalent system
approximations to aircraft response characteristics are then required for comparison with the quantitative
requirements.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.1.12.
The background information and user guide (BIUG) for that document, AFWAL-TR-81-3109, discusses
the rationale behind this requirement very succinctly, as follows:
In the past, both operational experience and flying qualities research were largely limited to aircraft which
behaved in the classical manner: response to control and disturbance inputs characterized by transfer
functions of familiar form. The effects of additional dynamics introduced through the flight control system
were recognized at the time MIL-F-8785B was written, but limited knowledge prevented adequate
treatment. Still, aircraft design developments continue to emphasize equalization to "improve" aircraft
response. In Systems Technology, Inc. TR-190-1, Stapleford discusses both good and bad possibilities.
Certainly one would expect that failure to consider one or more dynamic modes in the frequency range of
pilot control would give erroneous results. Prime examples include the F-14 (AGARD-CP-119) and the
YF-17 (AIAA Paper 75-985) designs. The F-14's stability augmentation system was designed to increase
the low short-period frequency. At one stage it did that well in landing approach, but it also introduced
higher-order dynamics which resulted in an overall effective short-period frequency little changed from
augmentation-off. In a flight evaluation of predicted YF-17 characteristics using the AFWAL-Calspan NT33 Variable Stability Airplane, pilots rated the short-period response poor to bad. The equivalent system
approach may not have been used to improve the response. However, it is pertinent that a configuration
intended to have good flying qualities got good pilot ratings in flight only after the flight control system
compensation had been simplified.
Boothe, et al. (AFFDL-TR-74-9) suggest several simple mechanizations that augment stability without
increasing the order of the system response. Nevertheless prefilters, forward-loop compensation,
crossfeeds, etc., are legitimate design tools that are being used on many current aircraft and indeed
seem to be the norm. These artifacts do increase system order and we need to be able to account for
their effects in the requirements. Thus, with modern flight control and stability augmentation systems,
there is considerable confusion regarding the proper selection of modal parameters such as short-period
frequency and damping. Correlation of Level 1 flying qualities with characteristics of the bare airframe is
certainly not valid for augmented aircraft in general. Stability and control augmentation frequently
introduce additional dynamics in the frequency range of pilot control, thereby invalidating any
interpretation of the requirements in terms
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APPENDIX A
of particular roots of a transfer function. Although these fallacies have been pointed out many times,
misinterpretations continue. The feeling is not uncommon that some requirements just do not apply. This
paragraph, 4.1.10, is intended to clarify application of the requirements to flying qualities in general.
In reality, we are only interested in pilots' opinions as to whether the actual aircraft dynamics enable the
appropriate tasks to be performed well enough with acceptable workload. We now require, therefore, that
the actual dynamics be approximated by the responses of transfer functions of classical form. The
appropriate parameters of this equivalent transfer function must meet the modal requirements of the
standard. This so-called equivalent system approach allows continued use of the familiar test data base
for a broad range of mechanizations. It has been advocated strongly by Hodgkinson and others (Systems
Technology, Inc. TR-190-1, AIAA Paper 77-1122, AIAA Paper 80-1627-CP, and Hodgkinson, Berger, and
Bear).
The preceding discussion should not be taken to imply that there is little problem with applying the
specification requirements to equivalent system parameters. For configurations which exhibit
conventional-appearing dynamics, application is indeed straightforward. It also appears to be true at
present that pilots are most comfortable with response dynamics that are natural, that is, like the classical
modes. Certainly, additional prominent modes result in a more complicated dynamic response. As we
consider configurations with dynamics that depart more and more from the classical order or form, then
more and more judgment will be required in defining the appropriate equivalent system parameters and
assessing compliance with the requirements. Hodgkinson has suggested that flying qualities will be poor
if no equivalent system can be found to give a good fit to the actual response. Although success of the
equivalent system approach in applying or defining the Level 2 and 3 boundaries is not definite at this
time, such application appears sufficient though possibly not necessary to achieve the desired goals.
There are also questions which remain to be answered. Is the equivalent system solution unique? (Not
universally, it seems.) Can the equivalent system parameters be juggled until compliance is indicated? (In
limited observations, some tendency toward equivalent results from different techniques has been noted.)
Are requirements necessary for either the amount or the quality of the mismatch? (To date this has not
been a major problem.) In spite of the qualifying remarks and the above questions, this approach is a way
to apply known requirements to advanced configurations with high-order dynamic responses. We
preserve the validated data base of MIL-F-8785B/C and the experience in its use. At the same time the
equivalent systems are to be defined by matching an appropriate aircraft response to pilot control input.
We therefore focus attention on the quality of the actual overall response perceived by the pilot, rather
than to imply consideration of a dominant mode as may be inferred (however incorrectly) from MIL-F8785B. We also believe that the use of the equivalent system approach is responsive to the needs of
designers. Failure of an equivalent system parameter to meet the requirement then indicates the
characteristic of the actual system (e.g., bandwidth, peak amplification, phase lag) that must be improved.
We acknowledge that the use of equivalent systems is not a magic solution to good flying qualities;
however, properly used it is a good tool for designing or evaluating advanced configurations which are
becoming indiscriminately complex.
In order to demonstrate compliance with the modal requirements of MIL-F-8785C, equivalent systems
must first be defined to approximate the actual aircraft dynamics whether predicted analytically or
obtained from flight test. Considerations for specific axes are discussed elsewhere following the
appropriate requirement. In general, however, it is necessary to add a term representing a time delay to
the "classical form" of the response. This term, itself a specified parameter (4.2.1.2, 4.5.1.5), allows a
closer match of the higher-frequency content of most advanced systems considered to date. The time
delay has been correlated with pilot opinion ratings.
The requirement as written is intended to allow the contractor to use any reasonable method of
determining the equivalent aircraft systems. However, the procuring activity may require some particular
method for final compliance demonstration. Guidance for some of the individual requirements contains
more detail.
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APPENDIX A
Other forms of dynamic requirements, such as time to bank, apply to the actual system, whatever its
order or nonlinearity. It is only the transfer-function parameters that are to be equivalent; whereas the
specified responses are generally those of the entire, actual system.
REQUIREMENT LESSONS LEARNED
In AFWAL-TR-82-3064, Powers notes that the large separation of short-period frequency and pitchnumerator inverse time constant makes the pitch transfer function for control inputs difficult to match with
the standard equivalent system. Whether this indicates a problem with equivalent systems or a flying
qualities deficiency is hard to determine. We do note that for low speed (with less separation of those
roots) NLR-TR-79127U reports rating degradation with the large pitch overshoot which results from the
separation. Some other requirements are suggested in this handbook for use when the equivalent system
approach is not satisfactory.
Hodgkinson and Bischoff (separately) discuss equivalent systems in AFWAL-TR-82-3064. AFWAL-TR81-3118 lists generally good matches of longitudinal short-term dynamics of some unstable aerodynamic
configurations stabilized by pitch rate feedback and forward-loop integration; pilot distance from the
center of gravity was found to be a complicating factor.
5.1.10 Interpretation of quantitative requirements-verification.
Verification shall be by analysis, simulation and test.
VERIFICATION RATIONALE (5.1.10)
Either frequency-response or time-response methods can be used (see Appendix B) to identify the
equivalent-system parameters. Appendix B outlines a procedure for matching the higher-order frequency,
response, however obtained. Frequency-response matching is preferred because it assures the best
match over the range of pilots' control-input frequencies in the closed-loop tasks of primary interest.
VERIFICATION GUIDANCE
Appendix B discusses matching techniques and a computer program for performing the match. Flight
measurement techniques for frequency response data are discussed in 5.2.1.2.
VERIFICATION LESSONS LEARNED
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APPENDIX A
4.1.11 General flying qualities requirements
4.1.11.1 Buffet.
Within the boundaries of the Operational Flight Envelope, there shall be no objectionable buffet which
might detract from the effectiveness of the aircraft in executing its intended missions. In the Permissible
Flight Envelope ___________.
REQUIREMENT RATIONALE (4.1.11.1)
The intent of this requirement is to prevent the occurrence of objectionable levels of buffet in the course of
operational flight.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.4.5.
"Objectionable" is to be interpreted in the context of operational missions: an annoyance, distraction or
discomfort; or an interference with maneuvering or tracking. For a combat aircraft the procuring activity
may need to extend the requirement to apply throughout the Permissible Flight Envelope. The extension
would apply, for example, to an air combat fighter intended to have high-angle-of-attack capability and to
a trainer for spinning.
Clearly, in those cases where buffet is a signal to the pilot of approach to a dangerous flight condition
(4.1.8.1) some buffet is desirable--but there should be no need for that within the Operational Flight
Envelope. AGARD-AR-82 contains a concise discussion on buffet and offers some guidelines on the
acceptability of various buffet levels:
To the fighter pilot who knows his aircraft, buffet onset is a valuable source of information
in moments of intense activity when he is not able to refer to his flight instruments. Of the
many different buffet level criteria to be found ... the following is a summary which
smoothes out the variations. The "g" values quoted are maximum excursions about trim:
Onset
+ .035 to .1 gz
perception depends on workload/normal g
Light
+ .1 to .2 gz
definitely perceptible
Moderate
+ .2 to .6 gz
annoying
Severe
+ .6 to 1.0 gz
intolerable for more than a few seconds
Provided that there are no other effects such as loss of full control or random aircraft
motions, light buffet usually had no adverse effect on maneuvering, either coarsely or
precisely. The average fighter pilot is so used to flying in this region that he may not even
comment on it at the lower amplitudes. He will however feel annoyance and frustration
when the buffet characteristics reach the level where his ability to track his target is
affected, other effects on his performance may result from the arm mass feedback to the
stick and his ability to see the target on his cockpit controls and instruments. At the
intolerable level the motion becomes physically punishing, and full control is not possible
as a result of the effect of the buffet on the pilot himself.
The significance of buffet in air combat depends upon the task. If flight in buffet gives a
performance improvement then pilots will use this region during the tactical phase of
combat. Tracking will also take place at quite high buffet levels, even with guns; but when
the low frequency, high amplitude "bouncing" buffet occurs then there is no further
advantage to be gained from operating in this region.
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APPENDIX A
REQUIREMENT LESSONS LEARNED
5.1.11 General flying qualities-verification
5.1.11.1 Buffet-verification.
Verification shall be by analysis, simulation and test.
VERIFICATION RATIONALE (5.1.11.1)
Flight testing at elevated angles of attack and load factors, and at lower angles transonically, will reveal
any buffeting tendencies. A windup turn maneuver while tracking a target can be especially useful in
identifying buffet regions. In flight, wing buffet intensity rise can be measured with a wingtip
accelerometer. Figure 9 illustrates methods of determining the region of buffet intensity rise from (a)
normalized rms values of wingtip normal acceleration and (b) estimations based on time history data.
VERIFICATION GUIDANCE
Wind-tunnel tests can give early indication of buffet onset and intensity, but flight testing will be needed to
determine the end effect with structural vibrations, noise, etc. included.
VERIFICATION LESSONS LEARNED
Judgment seems to be subjective, so marginal or critical cases should be evaluated by a number of pilots
138
MIL-STD-1797A
APPENDIX A
FIGURE 9. Buffet intensity rise determination from NASA-TP-1368.
139
MIL-STD-1797A
APPENDIX A
4.1.11.2 Release of stores.
The intentional release or ejection of any stores shall not result in objectionable flight characteristics or
impair tactical effectiveness of Levels 1 and 2. However, the intentional release or ejection of stores shall
never result in dangerous or intolerable flight characteristics. This requirement applies for all flight
conditions and store loadings at which normal or emergency release or ejection of the store is
permissible.
REQUIREMENT RATIONALE (4.1.11.2)
This requirement is included to insure that stores release will not have an adverse effect on flying
qualities.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.4.6 and 3.3.4.1.2.
Because of the variety of possibilities, this requirement must be left qualitative. All store loadings, internal
and external, which are specified in the contract are covered.
See MIL-HDBK-244 for additional requirements and guidance.
REQUIREMENT LESSONS LEARNED
5.1.11.2 Release of stores-verification.
Verification shall be by analysis, simulation and test.
VERIFICATION RATIONALE (5.1.11.2)
Evaluation of this criterion shall occur as a natural part of operational flight testing, usually preceded by
analysis (.e.g, AFFDL-TR-74-130 and AFWAL-TR-80-3032) and wind-tunnel testing. The wind-tunnel
tests may be guided on-line by trajectory calculations using a combination of currently-generated and
stored data.
There may be special flight envelopes in which store release or missile firing are permitted. Generally
such envelopes are ultimately cleared by flight testing.
VERIFICATION GUIDANCE
Store motion after release is such a function of the local airflow field that few generalities can be made
about the most critical conditions. At the same angle of attack the aerodynamic forces are greater at
higher speed. Dive angle, normal acceleration, store location, and store and ejector configuration and
loading are important to consider. The critical conditions will likely be at the boundaries of the release
envelopes.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.11.3 Effects of armament delivery and special equipment.
Operation of movable parts, such as bomb bay doors, cargo doors, armament pods, refueling devices
and rescue equipment, or firing of weapons, release of bombs, or delivery or pickup of cargo shall not
cause buffet, trim changes, or other characteristics which impair the tactical effectiveness of the aircraft
under any pertinent flight condition. These requirements shall be met for Levels 1 and 2.
REQUIREMENT RATIONALE (4.1-11.3)
This requirement is included to assure that armament delivery, etc., will not adversely affect flying
qualities, impairing mission effectiveness.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.4.7.
Because of the variety of possibilities, this requirement must be left qualitative. All armament and
equipment for the design missions are covered.
REQUIREMENT LESSONS LEARNED
Gun firing can cause deceleration and, depending on lateral and vertical location, attitude transients. It
has also been known to interfere with engine-inlet airflow or pilot vision. Rigidity and dynamics of local
structure and items attached to it influence the aircraft vibration resulting from gun firing.
5.1.11.3 Effects of armament delivery and special equipment-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5-1.11.3)
Operational flight test should be required, preceded by suitable analyses and wind-tunnel tests. Generally
the critical conditions should thus be known before flight test.
VERIFICATION GUIDANCE
Each movable part needs to be analyzed for its effect on stability and control.
VERIFICATION LESSONS LEARNED
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APPENDIX A
4.1.11.4 Failures.
No single failure of any component or system shall result in dangerous or intolerable flying qualities;
Special Failure States (4.1.7.2) are excepted. The crew member concerned shall be given immediate and
easily interpreted indications whenever failures occur that require or limit any flight crew action or
decision. The aircraft motions following sudden aircraft system or component failures shall be such that
dangerous conditions can be avoided by the pilot, without requiring unusual or abnormal corrective
action. A realistic time delay of at least between the failure and initiation of pilot corrective action shall be
incorporated when determining compliance. This time delay shall include an interval between the
occurrence of the failure and the occurrence of a cue such as acceleration, rate, displacement, or sound
that will definitely indicate to the pilot that a failure has occurred, plus an additional interval which
represents the time required for the pilot to diagnose the situation and initiate corrective action.
Additional requirements apply to transients from propulsion system (4.5.8.4, 4.5.9.5.5, 4.6.5.1, 4.6.6.2,
4.6.7.8) and flight control system (4.2.6.1, 4.2.8.6.5, 4.5.7.1, 4.5.9.5.6, 4.6.5.2, 4.6.7.9) failures.
REQUIREMENT RATIONALE (4.1.11.4)
These provisions involve safety of flight. In addition to accounting for flying qualities after a failure, we
recognize that the transient between the normal and the failed state could result in further flying qualities
degradation.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.4.8 and 3.4.9.
Recommended minimum time delay: A default value would be 1 second.
A minimum realistic time delay value of 1 second is consistent with Paragraph 3.3 9.3 in MIL-F-8785C.
For civil operation the FAA is more conservative with hardover failures of autopilot servos, requiring 3
seconds before pilot takeover is assumed. This time delay is to include an interval between the
occurrence of the failure and the occurrence of a cue such as acceleration, rate, or sound that will
definitely indicate to the pilot that a failure has occurred, plus an additional interval which represents the
time required for the pilot to diagnose the situation and initiate corrective action. The length of time should
correspond to the pilot's likely set to respond, for example longer during cruise than at takeoff.
NASA-CR-177331 or NASA-CR-177304 present guidance on determining a realistic time delay that
seems as applicable to winged aircraft as it is to rotorcraft. Table IX and the paragraph following it are
excerpts.
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MIL-STD-1797A
APPENDIX A
TABLE IX. Summary of minimum allowable intervention times for system failures.
Phase
Of
Flight
Attended
Operation
Divided
Attention
Operation
Hands On
Divided
Attention
Operation
Hands On
Unattended
Operation
Hands On
Unattended
Operation
Hands Off
Rotorcraft
Response
t1 - t0
Time for rotorcraft to
achieve change of
rate about any axis
of 3 deg/sec
OR
The time to reach
a change of "G"
in any axis of 0.2
OR
For an attention
getter to function
Time for rotorcraft to
achieve change of rate
about any axis of 3
deg/sec
OR
The time to reach a
change of "G" in any
axis of 0.2
OR
For an attention getter
to function
As above but the
threshold rates and "G”
values are
5 deg/sec and 0.25
respectively
Pilot
Response
t2 - tl
Minimum Allowable
Intervention Delay Time and
Method of Test
System failures will be
injected without warning to
the pilot. His ability to
recover as rapidly as
possible without a
dangerous situation
developing will be used to
assess system failure
mode acceptability.
1/2 sec
1-1/2 sec
(Decision 1
plus reaction 1/2)
2-1/2 sec (Decision
1-1/2 plus reaction
1)
The pilot will be warned of the
system failure. Demonstration
of compliance must show that
an intervention delay time
equal to 1 1/2 sec + (t1 – t0
can be tolerated
As above but intervention
delay time 2 1/2 seconds + (t1
– t0)
2-1/2 sec (Decision
2 plus reaction 1/2)
As above.
4 sec
(Decision 3
reaction 1)
As above but intervention
delay, time 4 sec + (tl – t0)
143
plus
MIL-STD-1797A
APPENDIX A
TABLE IX. Summary of minimum allowable intervention times for system failures - Continued.
ROTORCRAFT RESPONSE TIME INTERVAL (t1 - t0). This is the period between the failure occurring
and the pilot being alerted to it by a suitable cue. The cue may take the form of an adequate tactile, audio,
or visual warning. (The eye cannot be relied upon to distinguish abnormal instrument indications
sufficiently early for these to be regarded as an adequate cue). In the absence of the adequate cues
listed above, it can be assumed that a pilot will be alerted when the rotorcraft meets or exceeds the
responses listed for unattended operation.
PILOT RESPONSE TIME INTERVAL (t2 - t1). The period commences at the time the pilot is alerted to the
fact that something abnormal is happening and terminates when the controls are moved to commence
the recovery maneuver. The period consists of the recognition time, decision time, and reaction time. As
shown above, the recognition and decision times are assumed to increase as the pilot relaxes his level of
involvement, i.e., in going from "attended operation” to ”unattended operation” and also in going from
"hands on” to "hands off". The reaction time is longer "hands off" than ”hands on" as the pilot has to
locate the controls before he can move them.
*****************************
Pilot response time is especially critical in defining a reasonable minimum pilot
intervention delay time to a failure. The status of the pilot in the overall task of
controlling the rotorcraft can be described as active or attended control operation,
divided attention control operation (both hands on the controls and hands off), or
unattended control operation such as in autopilot mode (both hands on and hands off
the control). For example, if the pilot is making a final approach to a landing, he
would be considered to be in an attended operation mode of rotorcraft control with
his hands on the control. Should an automatic flight control occur, the minimum pilot
response time for corrective control input following recognition of the failure would be
quite small, approximately half a second. Therefore, for testing the acceptability of
failures in this mode of flight, it would be unreasonable to require testing (or
specification) of a minimum allowable response time any greater than 1/2 second.
However, for cross country flight at cruise airspeeds, it is very possible that the pilot
will not have his hands on the control if an autopilot is engaged. For failures which
have a significant probability of occurrence in this flight mode, the specification of a
1/2 second pilot response time for test purposes would be unreasonable and unsafe.
In this standard, therefore, the minimum allowable pilot response time would be
adjusted to 2-1/2 seconds following any single failure.
A propulsion failure along with a failed automatic compensation device or flight control system failure is a
consideration. A rational ground rule would be to include the probability of such combinations in the
calculations for 4.1.7.3 or 4.9.1 by assuming a probability of 1 that the critical engine failure occurs, and
adding any failures which result in nuisance actuation of the automatic device.
REQUIREMENT LESSONS LEARNED
Aircraft have been lost from runaway trim. That possibility needs careful consideration for every powered
trim system.
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APPENDIX A
5.1.11.4 Failures-verification.
Verification shall be by analysis, simulation, and test.
VERIFICATION RATIONALE (5.1.11.4)
For those failures and flight conditions judged too hazardous to evaluate in flight, demonstration likely will
be by simulation. Validated models of the aircraft and its flight control system will be needed for that, and
adequate motion cues should be available to simulate the acceleration environment with one-to-one
fidelity for at least two seconds following the failure.
VERIFICATION GUIDANCE
See table IX.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
APPENDIX A
4.1.11.5 Control margin.
Aerodynamic control power, control surface rate and hinge moment capability shall be sufficient to assure
safety throughout the combined range of all attainable angles of attack (both positive and negative) and
sideslip. This requirement applies to the prevention of loss of control and to recovery from any situation,
including deep stall trim conditions, for all maneuvering, including pertinent effects of factors such as pilot
strength, regions of control-surface-fixed instability, inertial coupling, fuel slosh, the influence of symmetric
and asymmetric stores (4.1.3), stall/post-stall/spin characteristics (4.8.4 through 4.8.4.3.2), atmospheric
disturbances (4.9.1) and Aircraft Failure States (4.1.7 through 4.1.7.6; failures transients and
maneuvering flight appropriate to the Failure State are to be included). Consideration shall be taken of the
degree of effectiveness and certainty of operation of limiters, c.g. control malfunction or mismanagement,
and transients from failures in the propulsion, flight control and other relevant systems. Additionally, for all
failure states and flight conditions, control margins shall be such that control can be maintained long
enough to fly out of atmospheric disturbances, all Flight Phases can be terminated safely, and a waveoff
(go-around) can be accomplished successfully.
REQUIREMENT RATIONALE (4.1.11.5)
This overall requirement is intended to assure adequate control for safety in any situation not otherwise
covered in the Standard. It is intended to permit recovery from unusual situations in, and even beyond,
the Permissible Flight Envelope--on the grounds that if a condition is attainable, someday it will be
attained. Experience has shown that to be a reasonable assumption.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.4.10.
To attain performance benefits, we no longer require control-surface-fixed stability. Whatever the cause,
control saturation can be catastrophic in a basically unstable airframe. Then control deflection for
recovery, whether commanded by the pilot or automatically, is just not available. This differs from the
stable case, in which if the deflection limit is reached for trim, full control authority is available for
recovery. Control rate limiting can also induce instability if the basic airframe is unstable. This
requirement, together with the related changes mentioned, is intended to require full consideration of all
the implications of relaxed static instability and other Control-Configured Vehicle (CCV) concepts.
In considering how much margin of control should be required there is no general quantitative answer, but
it is possible to enumerate some cases to consider. Certainly there should be sufficient control authority
to pitch the aircraft out of any trim point to lower the angle of attack from any attainable value. That is,
with full nose-down control the pitching moment should be at least a little negative at the most critical
attainable angle of attack, for a center of gravity on the aft limit and nominal trim setting. Attainable angle
of attack is another issue in itself; but lacking intolerable buffet or a limiter that is effective in every
conceivable situation, angles to at least 90 degrees should be considered. Control margin is also
necessary at negative angle of attack.
The flight task will dictate some minimum amount of nose-down control capability. Air combat
maneuvering certainly imposes such a requirement, and so do terminal-area operations including landing
flareout. Then, there should be some capability to counter atmospheric disturbances while maneuvering
and center-of-gravity movements due to fuel slosh while accelerating, diving or climbing; stop rotation at
the takeoff attitude, etc. Roll inertial coupling has been a critical factor for many slender aircraft
In addition to conventional control modes, a CCV's direct force controls can offer a number of new,
possibilities ranging from independent fuselage aiming to constant-attitude landing flares. The additional
variables must be accounted for to assure adequate sizing of the control surfaces, and priorities may
need to be established. The effectiveness of thrust vectoring varies with airspeed and altitude, and of
course with the commanded thrust level; engine flameout or stall may be a consideration.
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MIL-STD-1797A
APPENDIX A
The instabilities and complications resulting from these factors can probably be rectified by stability
augmentation if and only if control effectiveness is adequate. The controllability margin conventionally
provided by static stability must be translated for CCV's into margins of control authority and rate. Control
must be adequate for the combined tasks of trim (establishing the operating point), maneuvering,
stabilization (regulation against disturbances), and handling of failures (flight control system, propulsion,
etc).
4.1.12.8 precludes dangerous single failures. After the first failure it may be advisable to constrict flight
envelopes for some assurance of flight safety in case, say a second hydraulic system should fail. The
procuring activity will need to weigh the expected frequency and operational consequences of such
measures against predicted benefits.
Excessive stability, as well as excessive instability of the basic airframe, is of concern with respect to
available control authority and rate; for example large stable Clβ, increases the roll control power needed
to counteract gusts.
The requirements of this paragraph are largely an emphasis or amplification of other requirements in this
standard, among them:
4.1.1
Loadings
4.1.3
External stores
4.1.6.2
Flight outside the Service Flight Envelope
4.1.11.4
Failures
4.1.12.7
Transfer to alternate control modes
4.2.5
Pitch trim changes
4.2.7
Pitch axis control power
4.5.8
Roll axis control power
4.6.5.1
Yaw axis response to asymmetric thrust
4.6.6
Yaw axis control power
4.8.1
Cross-axis coupling in roll maneuvers
4.8.4
Flight at high angle of attack
4.9
Flying quality requirements in atmospheric disturbances
REQUIREMENT LESSONS LEARNED
It is well known that hinge moments can limit both deflection and rate of control surfaces. When using a
surface for control in two axes, as with a horizontal stabilizer deflected symmetrically for pitching and
differentially for rolling, priorities or combined limits must be set to assure safety (AIAA Paper 78-1500).
Other demands on the hydraulic system can reduce control capability at times. Aeroelasticity can reduce
control effectiveness directly, as well as alter the aircraft stability. For the F-16, full nose-down control put
in by stability augmentation has to be overridden in order to rock out of a locked-in deep stall. Also,
aerodynamics sometimes have to remind control analysts that control surfaces themselves stall at an
incidence somewhat less than 90 degrees; and if control is supplemented by thrust vectoring, for
example,
147
MIL-STD-1797A
APPENDIX A
one must consider the control force or moment available in normal operation, the effect on forward thrust,
and the possibility of flameout, as well as aerodynamic interference effects. All the possible interactions of
active control must be taken into account.
Encountering the wake vertex of another aircraft can be an extremely upsetting experience. These
encounters are not uncommon in practice or real air combat, and also may occur in the terminal area and
elsewhere; prediction is difficult. Other atmospheric disturbances can be severe, too: jet streams, storms,
wakes of buildings, etc., as well as gusts and wind shear.
The amount of control capability at extreme angles of attack, positive and negative, must be enough to
recover from situations that are not otherwise catastrophic. Avoidance of a locked-in deep stall has been
known to limit the allowable relaxation of static stability. Also, control must be sufficient to counter the
worst dynamic pitch-up tendency below stall or limit angle of attack. Propulsion and flight control system
failure transients must be considered, along with possibly degraded control authority and rate after failure:
spin/post-stall gyration susceptibility and characteristics may well be affected. Fuel system failure or
mismanagement must be allowed for.
The range of maneuvers considered should account for both the stress of combat and the range of
proficiency of service pilots. For example, in 1919 the British traced a number of losses of unstable
airplanes to control authority insufficient to complete a loop that had got flattened on top (ARC R&M No.
917). Thus nose-up capability at negative angles of attack can also be important. Poorly executed
maneuvers may make greater demands on the flight control system for departure prevention or recovery.
For CCVs as well as conventional aircraft, limiters can help greatly but their effectiveness and certainty of
operation need to be considered. Spins attained in the F-15 and F-16 attest to the possibility of defeating
limiters. AFWAL-TR-81-3116 describes the A-7 departure boundary's closing in with increasing sideslip
angle; angular rates also affect departure boundaries. Rapid rolling sometimes creates inertial coupling
which can put great demands on pitch control; nose-down pitching seems to accentuate the divergence
tendency.
External stores change both center of gravity and pitch moment (Cm0 and Cmα,). Experience with past
aircraft indicates a firm need to allow some margin to account for unforeseen store loadings. With relaxed
static stability this can determine not only the safety, but the possibility of flight with stores not considered
in the design process.
Uncertainties exist in the design stage. Nonlinear aerodynamics, particularly hard to predict even from
wind-tunnel tests, are almost certain to determine the critical conditions. The center of gravity (c.g.), too,
may not come out as desired. And in service the c.g. location is only known with limited accuracy. There
are also possible malfunctions and mismanagement in fuel usage to consider. We have even seen recent
cases (e.g., F-111 and F-16) of misleading wind-tunnel results on basic static stability. Aeroelasticity and
dynamic control effectiveness (e.g., F-15) can also reduce control margins.
Asymmetric loadings need to be considered. A critical case for the L-19 was the addition of a wire-laying
mission involving carriage of a large reel under one wing. Some aircraft - the F-15 is a recent example have been prone to develop significant fuel asymmetries due to prolonged, inadvertent small sideslipping
Dive pullouts (n > 1) will accentuate the effects of loading asymmetries. Some F-100s were lost from
asymmetric operation of leading-edge slats (nonpowered, aerodynamically operated on their own, without
pilot action), in dive-bombing pullouts.
Reconfigurable flight control systems add a new dimension to tracking and managing the available control
power.
The control margin requirements must be met with aerodynamic control power only, without the use of
other effectors such as thrust vectoring. This approach was chosen because experience to date with
current
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technology inlets and engines operating at the distortion levels typical of high angle of attack at low speed
dictates caution, due to the considerable uncertainty about reliability and dependability for use to stabilize
and control the vehicle. Throttle usage is also a factor. While this requirement does not preclude the
application of thrust vectoring for low-speed agility and super maneuverability performance
enhancements in the future, it does reinforce the position that current technology engines/inlets should
not be relied upon as the only means to assure flight safety, prevent loss of control or provide recovery
capability anywhere in the flight envelope. Should future technology advancements provide demonstrated
engine/inlet reliability at low speeds and high angles of attack, the procuring activity may allow this
requirement to be modified for multiple engine aircraft such that thrust vectoring from one engine out may
be used to meet it.
5.1.11.5 Control margin-verification.
Verification shall be by analysis, simulation and ground and flight test.
VERIFICATION RATIONALE (5.1.11.5)
This is a flight safety item. Analysis and simulation should precede or accompany careful buildup to
suspected critical flight conditions.
VERIFICATION GUIDANCE
We do not intend through flight demonstration in dangerous cases to show compliance with this
requirement. "The combined range of all attainable angles of attack and sideslip" may even extend
beyond the Permissible Flight Envelope, except for certain highly maneuverable fighter and trainer
aircraft. Flight test bounds will be established according to such requirements as MIL-S-83691. For
extreme flight conditions a combination of model testing--wind-tunnel, free-flight if necessary, and
hardware--and analysis will often be adequate. These extremes should be investigated in some way,
whether or not the aircraft incorporates a limiter. The scope of analysis, simulation and testing needs
careful consideration at the outset of a program. Then the progress must be monitored for possible
additional troubles.
AFWAL-TR-87-3018 gives guidance on determining control deflection and rate margins and calculating
the deflection-saturated departure boundary in the conceptual and preliminary design stages, based on a
reduced-order system with full state feedback. At high speed and high dynamic pressure, the system
bandwidth required is high, increasing the importance of high-frequency control-system modes, structural
modes and system noise amplification. At low speed and low dynamic pressure, design risks are related
to the limited ability of aerodynamic control surfaces to generate control moments. The lack of stabilizing
control moments beyond some angle of attack, or control-surface rate limits, will compromise transient
responsiveness. Describing-function analysis treats control limiting as a gain reduction, which in general
lessens the stabilizing effect of feedbacks. Statistically-based margins for gusts reduce, but do not
eliminate, the possibility of inadequate control margin.
Figure 10 indicates some critical parameters in the response of an unstable system to a step command.
Factors influencing some control-margin increments may be seen in table X. To these margins must be
added another nose-down control increment to counter the pitch-out tendency while rolling about the x
stability axis (flight path). As a first cut,
∆δ =
Ix − Iz
⋅ p 2 ⋅ sin 2α
2M δ I y
,
I − Iz
δ& = x
⋅ p ⋅ p& ⋅ sin 2α
MδIy
where p is the stability-axis roll rate (about the flight path). Figure 11 shows in concept the margins that
are needed: ∆δmarg is the sum of turbulence and sensor-noise components, ∆δtran provides the pitching
acceleration to meet the CAP requirement, and ∆δpr can cancel the inertial pitching moment from rolling.
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Unless deactivated whenever saturation is encountered, an integrator in the flight control system tends to
run away, leading to loss of control.
Similar considerations, also treated in AFWAL-TR-87-3018, apply to any basic airframe having static
directional instability.
VERIFICATION LESSONS LEARNED
While flight-test risk must be bounded, it is necessary to assure by some means that any dangerous
conditions are found and evaluated before service pilots and aircraft are lost through surprise encounters,
with no known avoidance or recovery technique. Flight experience can be summarized by Murphy's Law.
Therefore, it is better for highly skilled flight test pilots to find any serious glitches under controlled
conditions rather than to wait for some less experienced operational pilots to find them in service use.
During F/A-18 high α/stall testing, an α hang-up phenomenon was observed (at 50-60 deg α), which was
very similar to that described in 4.8.4.2.3 Lessons Learned with regard to the F-16 deep stall. At
operational aft c.g.s and high α, delayed recoveries were experienced in the F/A-18 due to weak nosedown pitch restoring moments, even with full forward stick. Based upon F/A-18 test experience a pitch
restoring moment coefficient (Cm) of at least -0.2 should be available for the most longitudinally unstable
loading/aft c.g. combination expected to exist on Class IV aircraft. Analysis of F/A-18 test data from high
α post-stall gyrations shows that the α hang-up phenomenon was further aggravated by uncommanded
roll rate and yaw rate oscillations and resultant nose-up pitching moments. Flight test results indicate that
these oscillations could generate pitching moments equivalent to approximately +0.1 ∆Cm, which
significantly opposed natural aerodynamic pitch restoring moments. Occasionally, F/A-18 recoveries from
high AOA hang-ups were significantly delayed because of accompanying roll/yaw rate oscillations when
c.g./loading/AOA conditions caused Cm (full nose down control input) to be less than approximately -0.2.
This suggests that for Class IV aircraft, a pitch recovery criterion could be that the pitch recovery control
produce a net pitch restoring moment of not less than -0.1 (approximately 15-20 deg/sec2 at low
airspeed).
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FIGURE 10. Control surface requirements.
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TABLE X. Control-margin increments.
Flying Quality
Stabilization
∆δFQ/∆nc = 57.3 CAP′/Mδ deg/g (for Teff ≤ 0.05)
Turbulence
σδ/σw function of Mw, Mδ, ωspcl, ζ spcl, structural modes
∆δstab / ∆n c = 57.3
g 1 / Tsp • 1 / Tsp
•
deg/g (linear, 2 DOF)
U0
Mδ • 1 / Tθ
1
2
2
Sensor noise
- most severe at low q
- 3σδ and σw for severe turbulence recommended
σδ/σs function of Ks, KF, ωs, 1/Ta, ω spcl, ζ spcl, ωsp ol
2
Flying Quality
δ& FQ /nc = 57.3 CAP/(Mδ • Teff) for desired CAP
Stabilization
δ& stab /nc < δ& FQ /nc if FCS stability margins OK & 1/Teff > ωc
δ& /nc function of 1/Teff, 1/Tsp , ωsp , ζ sp
stab
Turbulence
2
cl
σ δ& /σw function of 1/Ta, ωspcl, ζ spcl, Mδ
cl
- most severe at low q
- 3 σ δ& recommended for control margin
Sensor noise
2
σ δ& /σs = Ks KF • fn(ωs, 1/Ta and, for low ωspcl: ωspol , ωspcl, ζ spcl)
- these parameters are not all independent
- 3 σ δ& recommended for control margin
∆nc is the commanded increment of normal acceleration
1/T2 is the unstable pole of the transfer function (negative; 1/sec)
2
ωsp
is the 2-deg-of-freedom product of the poles, 1/sec2
ol
ωspcl and ζ spcl are the closed-loop frequency and damping ratio of the short-period mode
CAP is q& o /∆n∞, CAP′ is q& max / ∆n∞
ωs is the sensor bandwidth
Ks, KF are the sensor and forward-loop gains
σs, σw are the rms intensities of sensor noise and vertical gusts
& /nc transfer function
ωc is the crossover frequency of the δ
Teff is the effective time constant of command-path plus forward-path control-loop elements (such
as prefilters and actuators)
Ta is the time constant of the actuator ram
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FIGURE 11. Control margin requirements.
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4.1.11.6 Pilot-induced oscillations (PIO).
There shall be no tendency for pilot induced oscillations, that is, sustained or uncontrollable oscillations
resulting from efforts of the pilot to control the aircraft. More specific requirements are in 4.2.1.2, 4.2.2,
4.5.2 and 4.6.3. REQUIREMENT RATIONALE (4.1.11.6)
This general qualitative requirement, applicable to all axes, covers those axes of control for which there is
no database for more specific requirements.
REQUIREMENTS GUIDANCE
The applicable MIL-F-8785C requirements are paragraphs 3.2.2.3 and 3.3.3.
PIOs were a consideration in setting the boundaries of 4.2.1.2 and 4.5.1.3 through 4.5.1.5.
REQUIREMENT LESSONS LEARNED
Likely causes are equivalent time delay, control system friction, or inappropriately-located zeros of aircraft
transfer functions. See the discussion under 4.2.2 NORAIR Rpt No. NOR-64-143 discusses a number of
possible PIO mechanisms.
5.1.11.6 Pilot-induced oscillations (PIO)-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.1.11-6)
Tight tracking asks, aggressively performed, will be critical, Figure 12 describe, a PIO rating procedure
similar to the Cooper-Harper procedure of figure 6. Comparing the Level and rating descriptions, roughly
a PIO rating of 1 or 2 would be level 1, a 3 or 4 PIO rating level 2, a 5 PIO rating Level 3, and of course a
6 PIO rating extremely dangerous.
VERIFICATION GUIDANCE
Pilot-vehicle analysis in the manner described in the discussions of the cited paragraphs should help in
the design stage. Ground-based simulation may or may not show up any PIO tendencies. Flight
evaluation in variable stability aircraft is a valuable tool. Final determination will come from flight test of
the actual vehicle. PlOs have occurred in both pitch and roll.
VERIFICATION LESSONS LEARNED
Attention to flying qualities per se during flight control design will take care of many potential problems.
PlOs may occur early in the aircraft life as on the YF-16 high speed taxi test that got airborne before its
first flight, or later in service, as with the T-38 as more pilots got to fly it. If PIO is not found readily, it
should be sought during the flight test program.
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MIL-STD-1797A
APPENDIX A
FIGURE 12. PIO tendency classification from AFWAL-TR-81-3118.
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APPENDIX A
4.1.11.7 Residual oscillations.
Any sustained residual oscillations in calm air shall not interfere with the pilot's ability to perform the tasks
required in service use of the aircraft. More specific quantitative requirements are in 4.2.3.
REQUIREMENT RATIONALE (4.1.11.7)
This general qualitative requirement, applicable to all axes, covers those axes of control for which there is
no data base for more specific requirements.
REQUIREMENT GUIDANCE
The applicable MIL-F-8785C requirements are paragraphs 3.2.2.1.3 and 3.3.1.1.
Likely causes are flight control system nonlinearities such as valve friction or, especially in unpowered
flight control systems, control system friction or hinge-mounted nonlinearities.
REQUIREMENT LESSONS LEARNED
The X-29A, an unstable basic airplane, exhibits very noticeable control-surface activity during ground roll.
This is a result of a compromise which keeps the stability and control augmentation active on the ground
in order to assure flight safety in the event of bouncing or early lift-off.
5.1.11.7 Residual oscillations-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.1.11.7)
Any residual oscillations should become manifest during expansion of the flight envelope, if they have not
already been discovered during ground simulations with flight hardware.
VERIFICATION GUIDANCE
Flight or ground-roll conditions with high stability-augmentation gains may be critical.
VERIFICATION LESSONS LEARNED
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4.1.11.8 Control cross-coupling.
No controller shall create a secondary response which is objectionable (for Levels 1 and 2) or dangerous
(for Level 3). This requirement applies to all continuous and discrete controllers which affect the motion of
the aircraft.
REQUIREMENT RATIONALE (4.1.11.8)
To keep from needlessly increasing pilot workload, controllers should perform their functions without
deleterious side effects serious enough to cause objections for Levels 1 or 2, or a flight safety problem for
Level 3.
REQUIREMENT GUIDANCE
Applicable MIL-F-8785C requirements are paragraphs 3.4.3, 3.4.11, 3.6.2. 3.6.3 and 3.6.4.
This general requirement is in addition to specific requirements, among them the ones on pilot-induced
oscillations (4.1.11.6, 4.2.2, 4.5.2, 4.6.3), response to failures (4.1.11.4, 4.1.12.8 4.2.6.1, 4.5 7.1, 4.6.5.1,
4.6.5.2, 4.6.7.9), transients and trim changes, configuration or mode changes (4.1.12.7, 4.1.13 4, 4.2.5,
4.2.6.2, 4.5.7.2, 4.5.9.5.7, 4.6.5.3, 4.6.7.10), pitch/flight path/airspeed interactions (4.2.4, 4.3.1 through
4.3.1.2, 4.4.1, 4.4.1.1), lateral acceleration response to a roll command (4.5.4), roll control in steady
sideslips (4.5.5), yaw response to roll controller (4.6.2, 4.6.7.1), yaw axis control in go-around (4.6.7.7),
direct side force control (4.6.1.3, 4.7.1) and other general requirements of 4.1.11 - 4.1.13. Some example
applications are given below.
Operation of controllers intended for flight-path or speed control should not cause objectionable pitch
response characteristics, and vice versa. (It is recognized too that some coupling in the right direction
might actually be favorable. For example, a slight nose-up response to an increase in throttle would
improve the short-term flight path response for STOL aircraft operating on the back side of the powerrequired curve; but while using the throttle for airspeed control, pitching up is the wrong direction of
response to a thrust increase.)
Crossfeeds and feedbacks in the stability and control augmentation system should generally reduce the
severity of trim changes, if not eliminate them; but aerodynamic cross-coupling may vary greatly with
angle of attack.
“Unique Flight Characteristics of the AD-1 Oblique-Wing Research Airplane" discusses control cross
coupling problems encountered on NASA's first oblique-wing airplane. In particular a pitch-roll coupling,
was noted. For even a symmetric airplane, spoiler deflection for roll control may produce disconcerting
pitching; spoiler deflection as a direct lift control affects drag.
REQUIREMENT LESSONS LEARNED
AIAA Paper 78-1500 recounts a B-1 design problem with its rolling/pitching horizontal stabilizer. While
deflection limits can cause similar trouble, in this case limited hinge-moment capability at high speed
caused a roll command effectively to command pitch motion too.
5.1.11.8 Control cross-coupling-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.1.11.8)
Any difficulties should become apparent during normal maneuvering plus check of failure states and
control reconfigurations.
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VERIFICATION GUIDANCE
For the B-1 case cited, high dynamic pressure was critical. Other aerodynamic and inertial coupling is
often critical at high angle of attack, low dynamic pressure.
VERIFICATION LESSONS LEARNED
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MIL-STD-1797A
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4.1.12 General flight control system characteristics.
As used in this standard, the term flight control system includes the pitch, roll and yaw controls, direct
force controls including leading-edge and trailing-edge flaps, stability augmentation systems, trim
selectors and all mechanisms and devices that they operate, including the feel system. The requirements
of this section, which are directly related to flying qualities, are in addition to the applicable control system
design specification, for example MIL-F-87242 or MIL-C-18244. Some of the important mechanical
characteristics of control systems (including servo valves or actuators) are: friction and preload, lost
motion, flexibility, mass imbalance and inertia, nonlinear gearing, and rate limiting. Meeting separate
requirements on these items, however, will not necessarily ensure that the overall system will be
adequate; the mechanical characteristics must be compatible with the nonmechanical portions of the
control system and with the airframe dynamic characteristics.
4.1.12.1 Control centering and breakout forces.
Pitch, roll, yaw, direct lift and sideforce controls shall exhibit positive centering in flight at any normal trim
setting. Although absolute centering is not required, the combined effects of centering, breakout force,
stability and force gradient shall not produce objectionable flight characteristics, such as poor precisiontracking ability, or permit large departures from trim conditions with controls free. Requirements for
particular controllers are to be found in 4.2.8.5, 4.3.4, 4.5.9.4 and 4.6.7.11.
4.1.12.2 Cockpit control free play.
The free play in each cockpit control, that is, any motion of the cockpit control which does not move the
control surface in flight, shall not result in objectionable flight characteristics, particularly for smallamplitude control inputs.
4.1.12.3 Adjustable controls.
When a cockpit control is adjustable for pilot physical dimensions or comfort, the control forces defined in
3.4.4 refer to the mean adjustment. A force referred to any other adjustment shall not differ by more than
10 percent from the force referred to the mean adjustment.
REQUIREMENT RATIONALE (4.1.12 through 4.1.12.3)
Customarily, flight control system specifications refer to the flying qualities requirements. The overall
flying qualities requirements pertaining directly to the flight control system are placed here because of
their generality. They apply to all axes of control and response.
REQUIREMENT GUIDANCE
Related MIL-F-8785C paragraphs are 3.5.1, 3.5.2, 3.5.2.1, 3.5.2.2 and 3.5.2.4. Also, see the references
for 4.1.13.
Cockpit automation is becoming increasingly prominent. We are not yet in a position, however, to give
more than a little general guidance on division of duties and the interface between human and automatic
control. Time lags in displays will tend, of course, to be seen as effective delays in aircraft response.
It has been suggested that the crew be used for head-up, eyes-out tasks, situation awareness and
executive functions:
mission
tactics
environment
contingencies
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MIL-STD-1797A
APPENDIX A
While some functions that might be automated are:
mission planning
station-keeping
monitoring
rendezvous
alerting
navigation (including air drop)
routine functions
As exemplified by the Integrated Fire and Flight Control (IFFC) advanced development program, there
are cases in which partial automation is the best task solution. In that program the HUD displayed a box,
its size representing the range of operation of the IFFC system. The pilot would put this box on his air-toair target, and IFFC would put the pipper on the target. It was found helpful to leave roll control to the
pilot: he could easily track the target's bank angle, and keeping involved aided the pilot to take over
quickly after any failure in the high-bandwidth IFFC system.
These general qualitative requirements result from experience. Although the trim system is part of the
flight control system, for convenience special trim system requirements have been grouped separately
under 4.1.13.
A discernible neutral point (or trim or equilibrium point) should always be provided in manual pitch, roll,
yaw or direct lift and sideforce controllers. That is, if the pilot chooses to release a control, it should return
toward a neutral or trim state. If no cues are provided, the pilot will be forced to search manually for such
a trim condition. This can lead to poor maneuvering control or, in the extreme, to pilot-induced
oscillations. Allowable levels of friction may prevent absolute centering; that generally is alright. Throttles,
instead, generally have no feel but friction. That has been found acceptable for infrequent, trim-type
usage, but would cause difficulty if continuous control were attempted about a trim point with it, for one
thing, after moving the throttle the pilot has no trim-position reference.
Some small amount of free play may be desirable to prevent over sensitivity to unintended control
motions. However, in normal operations, and especially in high-demand times such as turbulence
penetration or air combat, free play can contribute to overcontrol, loss of accuracy and rapid pilot fatigue.
No numerical value has yet been found that appears generally adequate. The allowable free play would
seem to be a function of control-deflection sensitivity (angular acceleration per inch or degree of
movement) and possibly control-force sensitivity as well.
The ten percent force allowance for control adjustments is intended to assure keeping within the
threshold of pilot perception.
REQUIREMENT LESSONS LEARNED
5.1.12 General flight control system characteristics-verification.
Verification shall be by analysis and flight test.
5.1.12.1 Control centering and breakout forces-verification.
Verification shall be by analysis and test. Measurement of breakout forces on the ground will ordinarily
suffice in lieu of actual flight measurement, provided that qualitative agreement between ground
measurement and flight observation can be established
5.1.12.2 Cockpit control free play-verification.
Verification shall be by analysis and flight test.
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5.1.12.3 Adjustable controls-verification.
Verification shall be by inspection.
VERIFICATION RATIONALE (5.1.12 through 5.1.12.3)
Although simulation may be helpful for design purposes, final verification shall be of the flight hardware
installed in the aircraft.
VERIFICATION GUIDANCE
Compliance with qualitative requirements is to be assessed according to 4.1.9. Throughout the flight test
program, pilots should note any difficulty in meeting these qualitative requirements.
Except for assessing flight characteristics, of course, these verifications generally may be performed on
the ground.
Evaluation in flight should be made over the operational load factor and airspeed ranges, at the minimum
and maximum operational altitudes, and especially at high speeds, where required control surface
deflections are small. Operationally-oriented fine tracking will be a critical test.
VERIFICATION LESSONS LEARNED
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4.1.12.4 Rate of control displacement.
The ability of the aircraft to perform the operational maneuvers required of it shall not be limited by control
surface deflection rates in the atmospheric disturbances specified in 4.9.1. Control rates shall be
adequate to retain stabilization and control in the Severe disturbances of those sections. For powered or
boosted controls, the effect of engine speed and the duty cycle of both primary and secondary control
together with the pilot control techniques shall be included when establishing compliance with this
requirement.
4.1.12.5 Dynamic characteristics.
A linear or smoothly varying aircraft response to cockpit-control deflection and to control force shall be
provided for all amplitudes of control input. The response of the control surfaces in flight shall not lag the
cockpit-control force inputs by more than the angles specified, for frequencies equal to or less than the
frequencies specified: ____________________.
4.1.12.6 Damping.
All control system oscillations apparent to the pilot shall be well damped, unless they are of such an
amplitude, frequency and phasing that they do not result in objectionable oscillations of the cockpit
controls or the airframe on the ground, during flight and in atmospheric disturbances.
4.1.12.7 Transfer to alternate control modes.
The transient motions and trim changes resulting from the intentional engagement or disengagement of
any portion of the flight control system by the pilot shall be such that dangerous flying qualities never
result. Allowable transients are further specified in 4.2.6.2, 4.2.8.6.6, 4.5.7.2, 4.5.9.5.7, 4.6.5.3, and
4.6.7.10.
4.1.12.8 Flight control system failures.
The following events shall not cause dangerous or intolerable flying qualities:
a. Complete or partial loss of any function of the flight control system as a consequence of any
single failure (approved Aircraft Special Failure States excepted).
b. Failure-induced transient motions and trim changes either immediately upon failure or upon
subsequent transfer to alternate modes.
c. Configuration changes required or recommended following failure.
The crew member concerned shall be provided with immediate and easily interpreted indication whenever
failures occur that require or limit any flight crew action or decision. Allowable transients are specified by
axis 4.2.6.1, 4.5.7.1 and 4.6.5.2.
4.1.12.9 Augmentation systems.
Operation of stability augmentation and control augmentation systems and devices, including any
performance degradation due to saturation, shall not introduce any objectionable flight or ground handling
characteristics. Any performance degradation of stability and control augmentation systems due to
saturation of components, rate limiting, or surface deflections, shall be only momentary, and shall not
introduce any objectionable flight or ground handling characteristics. This requirement particularly applies
for all Normal States and Failure States in the atmospheric disturbances of 4.9.1 and 4.9.2 and during
maneuvering flight at the angle-of-attack, sideslip, and load factor limits of the Permissible Envelope. It
also applies to post-stall gyrations, spins, and recoveries with all systems, such as the hydraulic and
electrical systems, operating in the state that may result from the gyrations encountered.
4.1.12.10 Auxiliary dive recovery devices.
Operation of any auxiliary device intended solely for dive recovery shall always produce a positive
increment of normal acceleration, but the total normal load factor shall never exceed 0.8 nL, controls free.
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4.1.12.11 Direct force controllers.
Direct force controllers include direct lift control systems and lateral translation systems. Direct force
controllers which are separate from the attitude controllers shall have a direction of operation consistent
with the sense of the aircraft motion produced, by conveniently and accessibly located, comfortable to
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use and compatible with pilot force and motion capabilities. Transients encountered with engagement of
these modes shall meet the requirements of 4.1.12.7, 4.2.6.2, and 4.6.5.3. Functions should be provided
in the control system that would only allow this mode to be engaged within its design flight regime or
maneuvers. When used either by themselves or in combination with other control modes, flight safety and
mission effectiveness shall not be degraded. These systems shall not defeat limiters that are necessary
for stable and controlled flight, or for structural considerations.
REQUIREMENT RATIONALE (4.1.12.4 through 4.1.12.1 1)
This group of flying qualities requirements pertaining directly to the flight control system applies generally,
to all axes of control and response. The requirement on direct force controllers is written to ensure that
operation of the controllers is simple and straightforward. When implementing these controllers it must be
assumed that the pilot may elect to engage the device in the middle of a maneuver, or in conjunction with
another mode.
REQUIREMENT GUIDANCE
Related MIL-F-8785C paragraphs are 3.5.2.3, 3.5.3, 3.5.3.1, 3.5.6, 3.5.5, 3.4.9, 3.5.4, and 3.6.4.
These generally qualitative requirements, like the others in 4.1.13, result from experience. Compliance
with qualitative requirements is to be assessed according to 4.1.9.
Atmospheric disturbances in the form of gusts should not prevent any maneuvering in the Operational
Flight Envelope. This means that no limitations should be imposed due solely to control travel. Since
ability to counter gusts includes surface rate characteristics, these too are mentioned explicitly. While
specific disturbances are listed, the evaluation remains somewhat qualitative. The control required for
attitude regulation is in addition to that required for trim and maneuvering.
Auxiliary hydraulic devices may use up significant portions of the available hydraulic power during critical
phases of the mission. For example, actuation of landing gear, flaps, slats, etc., during the landing
approach when the engines are operating at relatively low power settings could drain enough hydraulic
power to make it difficult for the pilot to make a safe approach, especially in turbulence. In other flight
conditions with less auxiliary demand or higher engine thrust, that same hydraulic system might be more
than adequate. Also, at high dynamic pressure high hinge moments may limit control-surface rate and
deflection.
In precision control tasks such as landing approach and formation flying it has been observed that the
pilot sometimes resorts to elevator stick pumping to achieve better precision (see AFFDL-TR-65-198,
AFFDL-TR-66-2, and Boeing Report D6-10732 T/N). This technique is likely to be used when the shortperiod frequency is less than the minimum specified or if the phugoid is unstable, but has been observed
in other conditions also. Some important maneuvers, such as correcting an offset on final approach, call
for simultaneous, coordinated use of several controls.
In the Navy's experience the control surface lag requirements, which are not explicitly covered by the
flight control system specification, provide additional guidance concerning the portion of the time delay
which may be attributed to control surface lag. Such time delay can be an important source of pilotinduced oscillation tendencies.
4.1.12.5, if quantitative limits are desired on control surface lag, the limits from table XI are
recommended. While pilots do not normally observe surface motion due to stability augmentation, the
time delay between pilot input and surface movement (here expressed as a phase delay) can be
disconcerting--see 4.2 2 guidance.
164
MIL-STD-1797A
APPENDIX A
TABLE XI. Control surface lags.
Allowable Lag, def
Level
1
2
3
Category A & C
Flight Phases
15
30
60
Category B Flight
Phases
30
45
60
Control
Upper Frequency, rad/sec
Pitch
the large of equivalent ωsp
and 2.0
Roll &
Yaw
the largest of equivalent
ωd 1/TR and 2.0
The required operational maneuvers are commensurate with the particular level of flying qualities under
consideration. The maneuvers required in Level 3 operation, for example, will normally be less precise
and more gradual than for Level 1 and 2 operation. In some cases this may result in lower demands on
control authority and rates for Level 3 operation. Note, however, that when the handling characteristics of
the aircraft are near the Level 3 limits, increased control activity may occur, even though the maneuvers
are more gradual.
Another requirement on control-surface rate capability, more explicit although still qualitative, is 4.1.11 5.
Whereas that requirement concerns loss of control, 4.1.12.4 applies more generally.
“Dangerous flying qualities" need to be interpreted in a rational manner, so that it applies to feasible
design options. We cannot take all the danger out of flying or anything else.
The demands of various performance requirements and the rapid advancement of control system
technology have resulted in the application of relaxed static stability in both the pitch and yaw axes.
These systems provide excellent flying qualities until the limits of surface deflection or rate are reached.
In this case, the degradation in flying qualities is rapid and can result in loss of control due to pilot-induced
oscillations or divergence. It has been found, however, that momentarily reaching the rate of deflection
limit does not always result in loss of control; the time interval that a surface can remain on its rate or
deflection limit depends on the dynamic pressure, the level of instability of the vehicle, and other factors.
A thorough analysis of the capability of the augmentation system should be performed over the
Permissible Envelope and should include variations in predicted aerodynamic terms, e.g. position and
system tolerance. During flight at high angle of attack, operation of augmentation systems has caused
departure, either because the aerodynamic characteristics of the surface have changed or the surface
has reached its limit. During departures or spins, engines may flame out or have to be throttled back, or
shut down such that limited hydraulic or electrical power is available to control the gyrations, recover to
controlled flight, and restart the engine(s). The analysis of flying qualities should take into account these
degraded system capabilities.
If loss of control or structural damage could occur, an inhibit should be incorporated in the system such
that it cannot be engaged, or if it is already engaged, then other modes with which it is not compatible
cannot be engaged. Furthermore, operation of these devices should not be capable of defeating angle-ofattack limiters, sideslip limiters, or load factor limiters that are built into the basic flight control system to
provide stable and controlled flight.
REQUIREMENT LESSONS LEARNED
Following some failures, a pilot's lack of adaptation, or inappropriate adaptation, can result in a
pilot/airframe closed-loop instability, even if the aircraft itself remains stable. Examples are the B-58 yaw
damper and NASA-TN-D-1552.
165
MIL-STD-1797A
APPENDIX A
The required failure indications depend on operational rules. Consistent maintenance and checkout
capability and rules are needed.
On the X-29, SCAS gains were originally a function of air data from the side-mounted probe. Large
position error transonically gave erroneous gains which lowered the system's phase margin.
In certain flight conditions, turbulence intensities and failure states, performance or augmentation systems
can actually degrade the flying qualities. The purpose of this requirement is to ensure that this effect is
analyzed and minimized. Compliance with this paragraph is especially important to vehicles employing
relaxed static stability.
5.1.12.4 Rate of control displacement-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.5 Dynamic characteristics-verification.
Verification shall be by analysis and test.
5.1.12.6 Damping-verification.
Verification shall be by analysis and flight test.
5.1.12.7 Transfer to alternate control modes-verification.
Verification shall be by analysis and flight test.
5.1.12.8 Flight control system failures-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.9 Augmentation systems-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.10 Auxiliary dive recovery devices-verification.
Verification shall be by analysis, simulation and flight test.
5.1.12.11 Direct force controllers-verification.
Verification shall be by analysis, simulation, inspection, and flight test.
VERIFICATION RATIONALE (5.1.12.4 through 5.1.12.1 1)
Verification normally will be a part of flight control system design and testing, and of flight envelope
expansion.
VERIFICATION GUIDANCE
Evaluation pilots should be alert for potential operational problems in exploring the safe limits of the flight
envelope. Critical conditions will usually be the corners of the expected envelopes (e.g., a SAS for power
approach should be switched at the highest and lowest expected airspeeds, at low altitudes). Limited
analytical and ground-based simulation may be used to supplement actual flight testing, especially in the
early stages of development; but flight testing is ultimately required. The conditions examined should be
in the range of those encountered operationally.
For requirements involving flight in turbulence, compliance may be shown principally through analysis of
gust response characteristics using either an analytical model or a piloted simulation, involving the gust
models of 4.9 and a flight-validated model of the aircraft. Such analysis must include not only the normal
operational maneuvers involving pitch, roll, and yaw controls; but also the critical maneuvers (especially,
for hydraulic actuation systems) which may limit the responsiveness of the control surfaces. As mentioned
in Requirement Guidance, these might include extension of landing gear and high-lift devices on landing
approaches, etc. Some evaluation should be conducted by flying in real turbulence.
166
MIL-STD-1797A
APPENDIX A
A modicum of common sense is required in the application of this requirement. The specific intensities of
atmospheric disturbance to be applied are not specified. Yet section 4.9 contains turbulence up to the
thunderstorm level. We do not normally require operational maneuvering in thunderstorm turbulence. It
would seem reasonable to require operational maneuvering in turbulence intensities up to Moderate. For
turbulence intensities greater than Moderate it seems reasonable to require sufficient maneuver capability
for loose attitude control.
VERIFICATION LESSONS LEARNED
167
MIL-STD-1797A
APPENDIX A
4.1.13 General trim requirements
4.1.13.1 Trim system irreversibility.
All trimming devices shall maintain a given setting indefinitely unless changed by the pilot, or by a special
automatic interconnect (such as to the landing flaps), or by the operation of an augmentation device. If an
automatic interconnect or augmentation device is used in conjunction with a trim device, provision shall
be made to ensure the accurate return of the device to its initial trim position on removal of each
interconnect or augmentation command.
4.1.13.2 Rate of trim operation.
Trim devices shall operate rapidly enough to enable the pilot to maintain low control forces under
changing conditions normally encountered in service, yet not so rapidly as to cause oversensitivity or trim
precision difficulties under any conditions, including:
a. Dives and ground attack maneuvers required in normal service operation
b. Level-flight accelerations at maximum augmented thrust from 250 knots or VR/C - whichever is
less, to Vmax at any altitude when the aircraft is trimmed for level flight prior to initiation of the
maneuver.
4.1.13.3 Stalling of trim systems.
Stalling of a trim system due to aerodynamic loads during maneuvers shall not result in an unsafe
condition. Specifically, the entire trim system shall be capable of operating during the dive recoveries of
4.2.8.6.3 at any attainable, permissible n, at any possible position of the trimming device.
4.1.13.4 Transients and trim changes.
The transients and steady-state trim changes for normal operation of control devices such as throttle,
thrust reversers, flaps, slats, speed brakes, deceleration devices, dive recovery devices, wing sweep and
landing gear shall not impose excessive control forces to maintain the desired heading, altitude, attitude,
rate of climb, speed or load factor without use of the trimmer control. This requirement applies to all inflight configuration changes and combinations of changes made under service conditions, including the
effects of asymmetric operations such as unequal operation of landing gear, speed brakes, slats or flaps.
In no case shall there be any objectionable buffeting or oscillation caused by such devices. More specific
requirements on such control devices are contained in 4.2.5 and 4.1.12.10.
4.1.13.5 Trim for asymmetric thrust.
For all multi-engine aircraft, it shall be possible to trim the cockpit-control forces to zero in straight flight
with up to two engines inoperative following asymmetric loss of thrust from the most critical propulsive
factors (4.6.5.1). This requirement defines Level 1 in level-flight cruise at speeds from the maximumrange speed for the engine(s)-out configuration to the speed obtainable with normal rated thrust on the
functioning engine(s). Systems completely dependent on the failed engines shall also be considered
failed.
4.1.13.6 Automatic trim system.
Automatic trimming devices shall not degrade or inhibit the action of response limiters.
REQUIREMENT RATIONALE (4.1.13 through 4.1.13.6)
Establishing, maintaining and changing the trim or operating point are basic factors in piloting
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.6.1.4, 3.6.1.2, 3.6.1.3, 3.6.3 and 3.6.1.1
The trim system is part of the flight control system (4.1.12), but the general trim requirements are grouped
here for convenience.
The irreversibility requirement allows trim scheduling or interconnection with other control devices (e.g.,
flap) but it specifically disallows float or drift.
168
MIL-STD-1797A
APPENDIX A
It may be difficult to find a trim rate which will be good for all loadings, in all mission phases. Slow trim
rates will not keep up with rapidly changing flight conditions, and so will fatigue the pilot. Too rapid trim
rates give oversensitivity, make trim difficult and accentuate the effect of any runaway trim. 4.2.8.6.3 sets
specific limits on forces in dives and during rapid speed changes, while trimming to decrease the forces.
While the requirement on stalling of trim systems applies generally, the problem has been encountered
with pitch trim by adjusting incidence of the horizontal stabilizer. First, some of the available elevator
capability goes to oppose the mistrimmed stabilizer and less is left to counter any adverse gust-induced
pitching motions. Second, elevator forces will be increased and may complicate recovery from a highspeed dive. Third, and perhaps most significant, whenever the elevator opposes the stabilizer, the
aerodynamic hinge moment on the stabilizer may reach a level that is impossible for the trim actuator to
overcome. See, for example, AIAA Paper 64-353.
If, for example, nose-down trim is used to counter the aircraft's pitch-up response to a vertical downdraft,
the aircraft will pitch down more sharply when the draft reverses in direction. Elevator will be used to
counter the pitch-down motion, and the resulting aerodynamic load may be sufficient to stall the stabilizer
actuator when nose-up retrim is attempted. As speed increases, the adverse effects increase, and the
elevator may have insufficient effectiveness to counter the nose-down forces of the draft and the
mistrimmed stabilizer. It is obvious that tuck effects may also complicate the picture, and it is significant
that tuck effects cannot be countered by a Mach trim system that is unable to move the stabilizer.
In addition to requirements on the trim system, in 4.1.13.4 are limits on transients and control force
changes due to operation of other controls. Besides the pitch trim change requirements of 4.2.5, we
stipulate that no other control actions should add significantly to pilot workload.
REQUIREMENT LESSONS LEARNED
A Boeing 720B airliner encountered stalling of the pitch trim actuator during a turbulence upset over
O'Neill, Nebraska, on 12 July 1963 (NASA CR-2677). The aircraft was passing through 39,000 ft in a
climb to 41,000 ft in IMC when severe turbulence was encountered. A large downdraft was penetrated
and the aircraft pitch attitude increased to +60 deg. This occurred despite application of full forward stick.
The gust then reversed to a large updraft, putting the aircraft into a severe dive with an estimated flight
path angle of about -35 deg. The pitch trim control was reported by the crew to be frozen in the dive.
Recovery was made with power (pullout at 14,000 ft) and pitch trim control was restored.
Two other turbulence upsets occurred with commercial jet transports (another Boeing 720B and a DC-8),
in which the wreckage of both aircraft showed the trim actuator in the full nose-down position. The
frequency of such turbulence upset accidents has been reduced drastically in recent years by pilot
training to fly loose attitude control and to essentially ignore large airspeed excursions in severe
turbulence. However, the possibility of entering a dive with full nose-down mistrim should be considered
in the design process.
KC-135, B-57 and other aircraft have been lost due to runaway trim, so that now elaborate precautions
are commonly taken to preclude dangerous trim runaway, trim and control use of the same surface, or
trimming by adjusting the null position of the feel spring through a limited range. Civil airworthiness
regulations have long required ability to continue flight and land safely with maximum adverse trim.
Autotrim can be insidious. Several B-58s are thought to have been lost because the pitch autotrim would
allow approach to stall angle of attack with no indication whatsoever to the pilot until very close to stall.
Attitude-hold stabilization has a similar effect with the pilot's hand lightly on the control. Pitch autotrim
does not promote holding airspeed, and a number of trim and stabilization mechanizations need the
addition of some form of stall and overspeed limiters.
169
MIL-STD-1797A
APPENDIX A
5.1.13 General trim requirements-verification
5.1.13.1 Trim system irreversibility-verification.
Verification shall be by analysis and test.
5.1.13.2 Rate of trim operation-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.3 Stalling of trim systems-verification.
Verification shall be by analysis and flight test.
5.1.3.4 Transients and trim changes-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.5 Trim for asymmetric thrust-verification.
Verification shall be by analysis, simulation and flight test.
5.1.13.6 Automatic trim system-verification.
Verification shall be by analysis, simulation and flight test
VERIFICATION RATIONALE (5.1.13 through 5.1.13.6)
Most flight verification will be accomplished during the normal course of the flight test program.
VERIFICATION GUIDANCE
Flight test crews should monitor trim system characteristics throughout the program to note any
discrepancies. Of special interest are extreme loadings and corners of the flight envelope including
sustained maneuvers at n ≠ 1, e.g., dives and dive recoveries, pullups, wind-up turns, with the cockpit
trim setting fixed throughout--and for trim rate, rapid speed changes and configuration and thrust
changes. Check at the highest trim-system loadings, which may be the critical test of irreversibility.
It is clear that full nose-down mistrim should be accounted for in the dives. For example, a Boeing 720
with full nose-down trim at the dive entry will encounter stalling of the pitch trim drive in the dive if the pilot
is, manually attempting to pull out. Judgment will have to be applied to decide if the mission requirements
and failure considerations such as runaway trim or trim actuation power failure should allow this type of
abuse. See Requirement Lessons Learned for more discussion of this. FAA Advisory Circular 25.253-1A
gives guidance on design upset maneuvers for civil transport airplanes.
External stores can affect both stability and the zero-lift pitching moment as well as the aircraft loading.
VERIFICATION LESSONS LEARNED
170
MIL-STD-1797A
APPENDIX A
4.2 Flying qualities requirements for the pitch axis
Control force and deflection have both been found universally to be important pilot cues, so both controlsfixed and controls-free (control deflection and force) characteristics have been specified. Although control
of the flight path may be the pilot's ultimate aim, pitch attitude control is commonly used as a surrogate.
4.2.1 Pitch attitude dynamic response to pitch controller
4.2.1.1 Long-term pitch response.
Any oscillation with a period of 15 seconds or longer shall have the following damping:. Except as may be
provided in 4.4.1, 4.4.1.1 and 4.2.1.2, no aperiodic flight path divergence is allowed within the Service
Flight Envelope for any Level of flying qualities. These requirements apply with cockpit controls fixed and
with them free.
REQUIREMENT RATIONALE (4.2 through 4.2.1.1)
The long-term response, characteristically the phugoid mode involving airspeed and pitch attitude, is
important in unattended or divided-attention operation, as well as being the means through which speed
is regulated.
REQUIREMENT GUIDANCE
The applicable MIL-F-8785C requirement is paragraph 3.2.1.2. Recommended values, as in MIL-F8785C, are based on the data and analysis presented in AFFDL-TR-69-72:
Level 1
equivalent ζp > 0.04
Level 2
equivalent ζp > 0
Level 3
T2 > 55 seconds
The equivalent phugoid damping ratio, ζp. is to be determined from the three-degree-of-freedom
equivalent classical longitudinal response; i.e. the transfer function
θ
δ es
or
θ
Fes
=
(
) (s + 1 T ) e
[s + 2ζ ω s + ω ] [s + 2ζ ω s + ω ]
K θ s + 1 Tθ1
2
p
p
2
p
θ2
− τ es
2
sp
sp
2
sp
is matched to the actual frequency response to determine best-fit values of the equivalent-system
parameters. The Level 3 requirement, stated in terms of time to double amplitude, is to be checked
directly from the time response of the actual aircraft, for both nose-up and nose-down control impulses. In
practice this matching should seldom be necessary: either a single lightly-damped mode will stand out at
low frequency or no more than a glance will be needed to see that the requirement has been more than
met. In unaugmented aircraft the phugoid is the lightly damped oscillation at which this requirement is
aimed. If stability augmentation suppresses the phugoid, the requirement obviously has been met and
exceeded; then the only concern will be that the augmentation has not introduced other difficulties.
We might consider an additional Level 3 requirement, to cover the case of two unstable phugoid roots,
which has been encountered. Such a limit is probably a good idea, but data are insufficient to establish a
firm value.
171
MIL-STD-1797A
APPENDIX A
Two other factors have been observed to alter this mode: thrust offset and compressibility; a stability
derivative Mu can result from vertical placement of the thrust line relative to the center of mass; too much
Mu of either sign can be destabilizing and cause increased trim changes and gust response. The
transonic trim change or tuck tendency is restricted in 4.4.1.1. Supersonically the common, slight longterm instability from lift-curve slope decrease with increasing Mach number has not generally been a
noticeable flying qualities problem. However, at Mach 3 the XB-70 experienced a troublesome sensitivity
of flight path to pitch attitude, a different but perhaps related problem.
Simple approximations for classical phugoid frequency and damping (valid only at subsonic speeds, see
AFFDL-TR-65-218) involve only airspeed, L/D and the static and maneuver margins (Etkin, "Dynamics of
Atmospheric Flight"):
ω
2
p
g
=& 2  
V
2
 − C mα C N α

 N m − c.g.
2ζ p ω p
2

hn
g
 = 2  
 V  hm

g
=& 2  
V
2
gC
=& 2   D
 V  CL
in an exchange of kinetic and potential energy at constant angle of attack. Combinations of stability
augmentation, center-of-gravity variation with loading, thrust, and compressibility can cause an unstable
phugoid oscillation, its decomposition into two aperiodic roots (with a possible divergence) or a
restructuring which involves the short-period roots as well.
Although AFWAL-TR-83-3015 indicates a damping ratio somewhat less than 0.04 to be satisfactory for
the L-1011 transport, we have chosen to recommend the MIL-F-8785B/C values on the bases that (a)
they are derived from systematic flight evaluations and (b) with relaxed static stability seeming to become
the norm, stability augmentation should not frequently be required only in order to meet the phugoid
requirement. When employed, the augmentation should improve rather than degrade the long-term
response.
Note that for the entire data base, all other aircraft response modes are stable. Although little data exists
on multiple Level 3 flying qualities, some such combinations can be unflyable.
REQUIREMENT LESSONS LEARNED
While stability augmentation can easily improve the phugoid damping, the reverse has also been
observed AFWAL-TR-81-3118 reports simulation of two stability augmentation schemes for a basically
unstable airframe. Pitch rate feedback with forward-loop integration eliminated the phugoid mode. Angleof-attack feedback, however, at high gain actually reduced phugoid damping. Evidently the small angleof-attack contribution had improved the damping, so reducing the excursions affected it adversely.
Pitch attitude stabilization, or integration of pitch rate feedback, improves the phugoid damping but, at
least for high gain, restricts the maximum steady normal acceleration and makes dδ/dV, the static stability
indicator, zero. Even for terminal flight phases, where required maneuverability is not great, some
adjustments are helpful. Several investigators have found that pilots appreciate the further addition of a
turn coordination feature that eliminates the steady control force in coordinated turns, thus removing an
undesirable feature of such augmentation, the need to push forward to recover. Similarly, Calspan
(NASA-CR-172491) has found it beneficial to insert a 0.2 r/s washout so that the nose drops normally, as
speed bleeds off in a flare or landing. Without the washout, a nose-down correction to lose altitude
requires pushing on the stick--an unnatural action which pilots are reluctant to do near the ground, at
least until the learn the technique well. However, that same washout, it was found, would saturate the
Shuttle cockpit controller when the pilot pushes over to capture the glide slope.
172
MIL-STD-1797A
APPENDIX A
Also, since the effectiveness of attitude or integrated angular rate stabilization depends upon the lift-curve
slope (stabilization drives a transfer function's poles toward its zeros, and the zero at -1/Tθ1, generally
changes sign as stall is approached), angle-of-attack limiting or inhibition may be necessary
(WADC-TR-58-82) with pitch feedback:
1
Tθ1
=&
Z M

 1− δ u
Z X
M δZu
=& − X u + u u 
Zδ M w
Zw 
1−
MδZw

2g
V
=&






 T V ∂ (T W ) 
 cos(α + i t ) − γ
 W − 2
∂u 



 C 1 − T sin (α + i ) − C
L1 
t
D α C L1 

W


+
 CL Cm 

C N α 1 − δ α 

 Cm CN 

δ
α 

2g
V
CD
 0 + CL
 C L
 1
∂C D

−
2
 CL
 α ∂C L
  Z CL
δ
1 +  T
c
C
Mδ
 







 − sin(α + i t )  T − V ∂ (T W )  

 W 2 ∂u  







the latter approximation holding for CLα, >> CD, parabolic drag (where CDo, is the zero-lift drag) and
neglecting thrust effects and CLδ. Note that in terms of the low-frequency path zero, from Aircraft
Dynamics and Automatic Control:
1
Th1
=&
1
g Z u  1 − (Z δ M δ ) (M u Z u ) 


−
Tθ1 V Z w  1 − (Z δ M δ ) (M w Z w ) 
so that generally 1/Tθ1 approaches zero at some angle of attack above that for zero 1/Th1, or dγ/dV. These
transfer-function numerator approximations apply only to cases of stabilization and control through the
same moment effector.
5.2 Flying qualities requirements for the pitch axis-verification
5.2.1 Pitch attitude dynamic response to pitch controller-verification
5.2.1.1 Long-term pitch response-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2 through 5.2.1.1)
Testing specifically for long-term flight characteristics is time-consuming. However, most long-term
deficiencies will surface during attempts to stabilize at a new flight condition, for example establishing a
steady climb or turn, changing airspeed, or flaring for landing.
173
MIL-STD-1797A
APPENDIX A
VERIFICATION GUIDANCE
Some mechanizations of stability augmentation may greatly enhance phugoid damping, but at the same
time alter some steady-state flight characteristics: for example, holding a new pitch attitude after removal
of a pilot command. Test pilots will need to evaluate any unusual characteristics qualitatively (Supersonic
Transport Aeroplane Flying Qualities).
VERIFICATION LESSONS LEARNED
174
MIL-STD-1797A
APPENDIX A
4.2.1.2 Short-term pitch response.
The short-term pitch response shall meet the following requirements for control inputs of all magnitudes
that might be experienced in service use: ______________.
REQUIREMENT RATIONALE (4.2.1.2)
Pitch control of conventional aircraft is a vital element of flying qualities, both as a primary control axis (for
example, in pointing the aircraft during gunnery) and as an indirect way of controlling the aircraft flight
path (for example, in glide path control for landing).
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.1.12, 3.2.2.1, 3.2.2.1.1, 3.2.2.1.2, and 3.5.3.
Short-term is defined in terms of a frequency range or a time span. The short-term pitch response
characteristics are universally regarded to be extremely important--so important that controversy over
both the form and the substance of requirements has continued for some years. Although we have found
that the existing data base may be extended by means of the equivalent system concept to cover a
number of more elaborate stability augmentation applications, problems remain. Of course there are
bound to exist some configurations that just don't fit. But even among the ones that do, several questions
remain unresolved, including:
Effects of pilot location and blended direct lift control have been observed and need to be
accounted for
A continuing controversy over the merits of n/α vs 1/Τθ2= (g/V) (n/α) has not yet been resolved
Various forms of Bihrle's Control Anticipation Parameter (CAP) relating initial pitching
acceleration to steady-state normal acceleration have been tendered
Tolerable values of effective time delay need to be pinned down more fully
Task dependence needs to be explored more fully
There are those who prefer, or insist upon, a time-domain form of criteria. (In the world of linear
systems a duality exists between time and frequency responses, so the question becomes just
what details the requirements should emphasize.)
We feel that the ζsp, ωsp, n/α form of MIL-F-8785B/C not only fits the data but has demonstrated its
effectiveness for a number of highly augmented aircraft as well as for classical response. That, then, is
the approach normally to be preferred. Also, presented herein are statements in terms of equivalent ωsp
Tθ2 instead of ωsp /(n/α). A time-domain alternative, based on second-order response but usable directly
2
with some higher-order responses, is also given. For cases with no good equivalent system match,
frequency-response criteria involving Nichols charts or the bandwidth of the actual higher-order system
are presented. We hope that current research will clarify the outstanding issues.
OUTLINE OF 4.2.1.2 GUIDANCE
In the meantime we present a preferred form (CAP) followed by alternatives:
Page
A. CAP or ωsp /(n/α), ζsp, τθ
2
B. ωsp Tθ2, ζsp, τθ
172
190
175
MIL-STD-1797A
APPENDIX A
C. Transient peak ratio, rise time, effective delay
217
D. Bandwidth, time delay
225
E. Pilot-in-the-loop criteria
237
F. Drop back and Nichols-chart aircraft boundaries
244
We strongly suggest using several or all of these criteria in an aircraft design, although that probably
would not be practical in a specification.
The large amount of this guidance reflects the importance of short-term pitch response, the high attention
it has been given, and the great need for further study to derive a clear-cut, generally applicable set of
requirements.
Equivalent systems and their determination are discussed under the recommended form of short-term
requirement, A. CAP…, followed by various different interpretations of CAP, possible modifications, and
for Level 3 the allowance of a slight instability. Supporting data for CAP are deferred to B, ωsp Tθ2, since
the same data and similar interpretations are involved. The discussion groups the Level 1 and Level 2
supporting data for conventional aircraft by Flight Phase Category, according to the flight tasks which
were rated. For each Flight Phase, first ωsp /(n/α) and then ωsp Tθ2 is treated. The data base (Neal-Smith
2
and LAHOS) is discussed describing complications due to stability and control augmentation. The Level 3
requirements, repeated under each heading A - F, are based on other considerations, as discussed at the
end of Section A. Supporting data for Level 3 are presented at the end of Section B, under "Equivalent
System Data.”
Several other approaches are then given and discussed: Hoh's bandwidth criteria, Chalk's time-response
parameters, the Neal-Smith closed-loop criteria and Gibson's combination of time- and frequencyresponse rules. These criteria should be useful as additional design guidance, or one or more of them
can be used when the recommended form does not work. Supporting data are given for each form.
While normally force is considered the primary cue of pilot feel, both control force and control position are
significant. With a deflection control system, force feel will lag deflection because of the finite bandwidth
of the feel system. Where a pilot-force pickoff (say, a strain gauge) is used, any significant deflection of
the pilot control is effected through a follow-up. It has been found that a displacement controller should
not be penalized unduly for the feel-system lag, since pilots sense both stick input and stick output and
can compensate. The response requirements are to be applied to pilot control inputs accordingly.
A. CAP or ωsp /(n/α), ζsp, τθ
2
The recommended requirement is:
The equivalent pitch rate and normal load factor transfer-function forms for pilot control-deflection or
control-force inputs
(
) (s + 1 T ) e
=
[s + 2ζ ω s + ω ] [s + 2ζ ω s + ω ]
K s (s + 1 T ) e
n ′ (s )
=
or F (s) [s + 2ζ ω s + ω ] [s + 2ζ ω s + ω ]
K θ s s + 1 Tθ1
q (s )
δ es or Fes (s)
2
p
n
z
δ es
2
p
p
2
es
p
p
2
p
− τes
θ2
2
sp
h1
2
2
sp
sp
− τns
sp
sp
2
sp
simultaneously shall be fit to the corresponding actual response of the aircraft over a frequency range of
0.1 to 10 radians per second. The parameter n ′z is normal acceleration at the instantaneous center of
176
MIL-STD-1797A
APPENDIX A
rotation for pilot pitch control inputs. Corresponding matches shall be made for pilot control-deflection
inputs. The requirements apply to the equivalent-system parameters determined from the best match for
force inputs, and also for deflection inputs; the procuring activity will be the judge of the adequacy of the
matches. The parameter CAP is to be estimated from the equivalent-system parameters l/Tθ2 and ωsp
[in ft − lb − rad − sec units; n α =& ( V g) (1 Tθ2 )] or alternatively
(
CAP = ω2sp q& maxHOS
) [(n α ) (q& maxLOES )]
where HOS refers to the actual higher-order system and LOES refers to the equivalent lower-order
system; q& max is the peak value of pitching acceleration for a step input of pilot force or deflection.
In addition to the requirements of figure 13, for Category C Flight Phases, ωsp and n/α shall be at least:
CLASS
min ωsp
I, II-C, IV
II-L, III
0.87r/s
.7
LEVEL 1
min n/α
2.7g/rad
2.0
min ωsp
LEVEL 2
min n/α
0.6r/s
0.4
1.8g/rad
1.0
For Level 3, T2, the time to double amplitude based on the value of the unstable root, shall be no less
than 6 seconds. In the presence of any other Level 3 flying qualities, ωsp shall be at least 0.05 unless flight
safety is otherwise demonstrated to the satisfaction of the procuring activity. T2 applies to the value of an
unstable first-order root: T2 = -( l n 2)/λ where λ is the value of the unstable root.
Requirements on the equivalent pitch time delay, τθ , apply to the value for θ(s)/δes(s) for a deflection
control system (pilot controller deflection commands the control effectors) and to θ(s)/Fes(s) for a force
control system (pilot controller force commands the control effectors):
LEVEL
ALLOWABLE TIME DELAY
1
0.10 sec
2
0.20
3
0.25
In the event that an adequate match cannot be found, the contractor with the concurrence of the
procuring activity, shall substitute an alternative requirement.
177
MIL-STD-1797A
APPENDIX A
FIGURE 13. Short-period dynamic requirements.
178
MIL-STD-1797A
APPENDIX A
THE EQUIVALENT SYSTEMS CONCEPT
The advent of high-order feedback control systems has required specification of flying qualities
parameters for systems of, for example, twentieth order or more. Some of the additional modes may no
longer be well separated in frequency from 1/Tθ1, 1/ Tθ2, ωp and ωsp. In fact there are commonly a number
of high-order system characteristic roots in the short-period frequency range.
One approach to reducing the order of these systems for specification purposes would be to choose a
suitable short-period pair of roots from the high-order array. This was done either by picking dominant
roots (the most suitable pair) or by root tracking (identifying the locus of aircraft short-period roots as the
various feedback pins were increased). Considerable experience in recent years indicates that such an
approach will lead to unsatisfactory handling qualities. A clear responsibility of the standard is therefore to
discourage that particular type of order reduction. Considerable research has been devoted to order
reduction by matching frequency responses to obtain low-order equivalent systems. This concept,
originally developed more than ten years ago (AFFDL-TR-70-74 and Systems Technology Inc., TR- 1901), has been refined considerably (for example, see AIAA Paper 77-1122, AIAA Paper 79-1783, MDC Rpt
A6792, AIAA Paper 80-1611-CP, AIAA Paper 80-1627-CP, "Analysis of High Order Aircraft/Flight Control
System Dynamics Using an Equivalent System Approach", “Handling Qualities of Aircraft with Stability
and Control Augmentation Systems -- A Fundamental Approach," “Equivalent Systems Approach for
Flying Qualities Specification," and MDC Rpt A5596). Since this approach has the advantage of including
the effects of adjacent modes, is easily related to previous specifications, and identifies important delay
effects, it is recommended. In AFWAL-TR-82-3064 Gentry describes matching techniques, Hodgkinson
and Wood discuss the effects of mismatch and Bischoff shows how to allow for the higher-order system
effect of reducing the maximum pitching acceleration for a step input.
The use of lower-order equivalent systems allows us to extend to many higher-order systems the
application of well-established boundaries generated from classical airplane data, i.e., MIL-F-8785C.
More specifically, requirements for pitch axis control have set boundaries on the classical modal
parameters in1
θ(s)
δ(s)
=
(
) (1 Tθ2 ) e −τθ s
[ζ p ; ωp ] [ζ sp ; ωsp ]
M δ 1 Tθ1
This expression is a linearized, reduced-order model of the actual aircraft response. In most cases the
phugoid and short period modes are sufficiently separated that further order reduction is possible as
follows:
(
M δ Z w 1 Tθ1
[
M α ζ p ; ωp
]
)
and
Phugoid
(
[
)
M δ 1 Tθ2 e − τθs
s ζ sp ; ω sp
]
Short Period
These expressions, universally recognized as pitch models of phugoid and short-period dynamics within
appropriate respective frequency ranges, may normally be used in place of the three-degree-of-freedom
pitch equivalent system specified; likewise with normal load factor.
1
In this shorthand notation (1/T) represents (s + 1/T) and [ζ; ω] represents [s2 + 2ζωs + ω2]
179
MIL-STD-1797A
APPENDIX A
While no specific guidance on the lower frequency bound of the matching region is offered, the phugoid
and short period are generally separated by at least a factor of 10, which should be adequate to consider
them separately. The assumption of widely separated phugoid and short period modes breaks down at
low values of static stability (i.e.,Mα = 0) such as for conventional aircraft with extreme aft center of gravity
locations and on most STOL configurations. (Transonic tuck occurs when a nose-down pitching moment
with increasing Mach number causes the phugoid poles to split into two real roots, which may become
large. That is a separate issue, covered in 4.4.1.1.) In this region of near-neutral stability, pilot ratings
seem to be insensitive to the exact root locations and are Level 2 as long as damping is sufficient
(AFWAL-TR-82-3014, AFFDL-TR-72-143, etc.). In that case Schuler's criteria (AFWAL-TR-82-3014) are
the best available, though they are not given here because they are not based on flight experience. A key
issue during the development of the lower-order equivalent system (LOES) approach was whether to fix
or free 1/Tθ2 during the fitting process. When 1/Tθ2 is allowed to be free it can take on very large (or small)
values. If its physical significance were related purely to attitude control, it would be appropriate to utilize
the freed value when making comparisons with the criterion boundaries. However, considerable evidence
indicates that the role of 1/Tθ2 in the correlations of classical aircraft is more related to the lag from the
response in attitude to the response in path:
γ( s )
θ(s)
NFγ
∆
= es
NFθ
∆
es
=&
1
Tθ2 s + 1
Then the ωsp /(n/α) boundaries would be interpreted as a specification on path as well as attitude control.
2
The appropriate value of n/α to plot on the criterion would therefore be the fixed (real) one: with a single
control surface (e.g., no DLC), simple block-diagram algebra shows that no augmentation can change tile
dynamic relationship of pitch to heave motion.
An example of the differences with 1/Tθ2 fixed and free is seen in table XII (taken from MDC Rpt A6792
fits of the AFFDL-TR-70-74 data). It can be seen that substantial differences in all the effective
parameters exist between the 1/Tθ2 -fixed and -free fits. Hence the dilemma is not a trivial one.
TABLE XII. Examples of variations in LOES parameters with 1/Tθ2 fixed and free.
CONFIGURATION
1A
FIXED
1.25
1G
1.25
2H
1.25
4D
1.25
1/Te
FREE
0.43
176.
FIXED
3.14
ωe
FREE
2.54
ζe
τe
FIXED
0.39
FREE
0.65
FIXED
0
FREE
0.020
0.78
1.55
0.74
1.07
0.185
0.043
4.08
2.55
3.80
0.80
0.52
0.126
0.098
5.25
3.47
4.61
0.58
0.23
0.169
0.111
In general, the lower-order approximation will always retain the proper relationship between attitude and
flight path if the pitch-rate and normal-acceleration transfer functions are matched simultaneously.
Although the importance of pilot location to the assessment of flying qualities has been demonstrated
(e.g., AFWAL-TR-81-3118), at present our understanding is insufficient to account for it. Therefore in
order to
180
MIL-STD-1797A
APPENDIX A
avoid the introduction of more extraneous parameters we use n′z , the normal acceleration at the
instantaneous center of rotation (at xcr = Zδ/Mδ the initial nz response to a step control surface input is
& = q).
zero. It is assumed that measurements of θ and n′z are in level, symmetrical flight, so that θ
Conventional subscripts on the equivalent dynamics are retained for consistency; equivalent is implicit.
Normally, for short-term pitch control the numerator of the normal-acceleration transfer function has two
high-frequency factors which can be either complex or real, depending upon the point chosen along the
body: az = &z& - Ix &θ& . A point can be found (the instantaneous center of rotation) at which the two factors
are extremely high frequency and so can be ignored, leaving the az /δes numerator a constant (in 2
degrees of freedom) for most aircraft. The technique of simultaneously matching pitch and center-ofrotation load-factor responses is based on this supposition. [Actually the classical nz response at the
instantaneous center of rotation is only approximated with this simplified numerator; the generally small
difference is taken up in the time delay, τn, which is not forced to equal τθ.] It guarantees the same
frequency and damping ratio in both responses, as well as a dynamic relation of flight path to pitch rate
conforming to kinematic and physical principles, without introducing unnecessary new parameters.
Allowing τn to be free should partially account for the neglected zeros and any other differences.
However, we have come across two exceptions to the validity of this constant-numerator approximation
for the short-term nz response. Poles and zeros added by the stability and control augmentation may
cancel in some transfer functions but not others. Also, multiple surfaces may not all have the same
dynamics -- for example a washed-out direct-lift nap tied to the pitch controller. For these cases we
recommend using an equivalent (n/α)e obtained from the equivalent Tθ2:
(n α )e = (V g) (1 Tθ2 )e
and using good judgment in assessing the validity of the match. Direct-lift control effects on flying qualities
are treated further in 4.3, to the extent that present knowledge will permit.
This usage is the result of lengthy discussions about the significance of 1/Tθ2 . Proponents of holding
1/Tθ2 fixed argued that n/α in the specification is centered about path control. Holding 1/Tθ2 fixed at the
value determined from the lift curve slope, on the other hand, preserves the known path to attitude
relationship given above. Free 1/Tθ2 tends to gallop to large values for aircraft with known deficiencies,
thereby, revealing the existence of a problem. For example, a current high-performance fighter is known
to be rated excessively sluggish (Level 2) in the power approach flight condition. Figure 14 shows the
characteristics with 1/Tθ2 fixed and free. For 1/Tθ2 fixed, the sluggish response is manifested as excessive
effective time delay (0.15 sec) whereas for 1/Tθ2 free the deficiency is manifested as an n/α which falls on
the lower specification boundary. Finally, utilizing the 1/ Tθ2 -free fit to ωsp but plotting the fixed value of
n/α actually predicts an airplane with excessive abruptness (plots above the upper limit in figure 14). In
this case, as in most such instances, either method predicts the same Level of flying qualities but
manifests the causes very differently.
It should be noted that a perfect fit using both the attitude and flight path transfer functions will always (for
a single pitch control surface) yield the fixed value of 1/Tθ2. However, if there are lags, such as from a
stick prefilter, introduced at frequencies in the middle of the fitting region, the fit may be marginal:
whereas the lower-order equivalent system (LOES) of the θ/Fes transfer function is of first over second
order, with a first-order low-pass prefilter a good-fitting lower-order system turns out of form a first- over
third-order transfer function.
181
MIL-STD-1797A
APPENDIX A
τe values shown are equivalent delays in seconds
M is mismatch, 20/n
∑ (∆G
2
db
2
+ .02 ∆Pdeg
)
FIGURE 14. Effect of fitting with 1/ Tθ2 fixed and free, Category C requirements.
182
MIL-STD-1797A
APPENDIX A
The problem can be approached in two ways: 1) we can ignore the mismatch and use the LOES model;
or 2) we can utilize LOES modes more appropriate to the controlled element rather than being
constrained to a first- over second- order. The problems with the second alternative are that our data
base is for classical unaugmented airplanes, and that the requirements would vary with mechanization of
the flight control system -- a concept we prefer to avoid. The consequence of the first alternative is that
the fitting routine could come up with parameter values which are not physically meaningful. The
consensus was to accept the mismatch (Alternative No. 1), rather than attempt to expand the criterion.
DETERMINATION OF EQUIVALENT SYSTEMS; MISMATCH
The equivalent lower-order parameters for this section may be obtained by any means mutually
agreeable to the procuring agency and contractor. The equivalent system matching routine outlined in
Appendix A is provided as guidance to indicate the expected level of rigor in the matching procedure. The
representation specified in the requirement is not intended to require complex denominator roots, i.e. the
denominator may have two first-order roots rather than an oscillatory pair. In this case the short-period
roots are, with ζsp > 1:
1/T1 , 1/T2 = ζω ± ω ζ − 1
2
The parameters of the specified equations should be obtained by matching the high-order pitch response
and the normal load factor response from ω1 to ω2 with the frequencies defined as follows:
ω1
Normally 0.1 rad/sec but > ωp
ω2
Normally 10 rad/sec, but > the resulting equivalent ωsp and 1/Tθ2
Thus iteration may be necessary in uncommon cases. The purpose is to assure that the dynamics of the
equivalent airframe are adequately defined, without requiring unusually low- or high-frequency end points
in the match. Pilot control inputs up to 20 r/s have been observed, but the cutoff frequency for effective
piloted control of highly maneuverable aircraft seems to be slightly less than 10 r/s.
There is currently insufficient data to place definitive requirements on mismatch between the HOS and
LOES. It should be noted, however, that the question of mismatch is inherent in any n-dimensional
specification of an m-dimensional response, when n < m (Hodgkinson in AGARD-CP-333). For equal
weighting at all frequencies, mismatch is defined as:
M = Σ (∆G)2 + KΣ (∆φ)2
= Σ(GHOS - GLOES)2 + KΣ(φHOS - φLOES )2
where G is the amplitude in dB and φ is the phase in radians. ∆G and ∆φ are calculated at discrete
frequencies between ω1 and ω2 evenly spaced on a logarithmic scale and may be compared with the
envelopes in figure 15. The significance of a given frequency can be judged by the latitude of match
allowed.
A brief NT-33 landing approach simulation tackled the question of mismatch (AIAA Paper 79-1783 and
AFWAL-TR-81-3116). High-order systems and simulations of their low-order equivalents were flown. The
experiment indicated that very large mismatches proved unnoticeable to pilots (a sum-of-squares
mismatch around 200 in the frequency range of 0.1 < w < 10 rad/sec compared to the previous arbitrary
limit of 10).
MDC Rpt A6792 offers a theory to explain the adequacy of such large mismatches. By examining pilot
rating differences between pairs of configurations in previous NT-33 data (AFFDL-TR-78-122 and
183
MIL-STD-1797A
APPENDIX A
AFFDL-TR-70-74), frequency response envelopes were derived. Each pair of configurations consisted of
an unaugmented, low-order aircraft response and a high-order system formed by adding terms to the loworder response; figure 15 shows the envelopes that were derived. Some rough guidance is available from
these envelopes, which are an approximate measure of maximum unnoticeable added dynamics.2
As would be expected, the pilots were most sensitive to changes in the dynamics in the region of
crossover (1 - 4 rad/sec). Mismatches between the HOS and LOES in excess of the values shown in the
figure 15 envelopes would be cause to suspect that the equivalent parameters may not accurately predict
pilot opinion. In such cases it is recommended that criteria applicable directly to the actual system, such
as some of those in following section of this guidance, be used in place of the equivalent system form.
Additional comments on the use of equivalent systems may be found in AGARD-CP-260 and "Low-Order
Approaches to High-Order Systems: Problems and Promises". The influence of mid-frequency added
dynamics on LOES was discussed in “Low-Order Approaches to High-Order Systems: Problems and
Promises", where it was shown that a series of (possibly unrealistic) lead/lag combinations evaluated in
the Neal-Smith inflight simulation (AFFDL-TR-70-74) produce LOES parameters which are not
necessarily equivalent to their classical counterparts. Of the ten configurations with added lead/lag
dynamics, only five are predicted accurately. For the five that failed, the equivalent dynamics (ζe, ωe, τe)
were predicted to have Level 1 flying qualities but were rated Level 2 by the pilots. Table XIII lists the
dynamics of the HOS and LOES for these configurations. With the exception of Configuration 1C, all have
ζe < 0.5 (though still greater than 0.35). Three have τe = 0. All but 1C have a first-order lag near the shortperiod frequency; 1C has a first-order lead near ωsp.
2
The basic aircraft dynamics were modified via equalization networks. Modifications that resulted in 1
pilot rating change were defined as maximum unnoticeable added dynamics.
184
MIL-STD-1797A
APPENDIX A
FIGURE 15. Envelopes of maximum unnoticeable added dynamics (AFWAL-TR-82-3064).
185
MIL-STD-1797A
APPENDIX A
TABLE XIII. Lead/lag configurations with Level 1 LOES and Level 2 pilot ratings.
CONF
1/Tθ2
1A
1C
2A
2B
7A
NOTES:
1.25
1.25
1.25
1.25
2.5
ζsp
0.69
0.69
0.70
0.70
0.79
HOS
1/T1
ωsp
1/T2
ω3
2.2
2.2
4.9
4.9
7.3
2.0
5.0
5.0
5.0
8.0
63
16
63
16
63
0.5
2.0
2.0
2.0
3.3
ζe
LOES
ωe
τe
0.39
0.67
0.46
0.42
0.44
3.14
3.02
5.96
5.67
8.23
0
.079
0
.059
0
RATING
PILOT
M
6,4
3,5,5
4.5
6,6
4,5
W
5
4
4
4,5
2
1. HOS from Neal-Smith (AFFDL-TR-70-74); LOES from MCAIR (MDC Rpt A6792)
2. Equivalent dynamics are Level 1 on MIL-F-8785C limits.
3. T1 is an added lead; T2 a lag;
3 an actuator.
Figures 16 and 17 show the effects of the added lead/lag combinations on these configurations. The net
effect is an apparent hump around ωsp characterized in the LOES match by a low equivalent damping
ratio (table XIII). The lower-order form has no other way to match a hump in the amplitude plot. Similar
effects are seen in the phase angle, figure 17: the humps appear as phase lead (since, for the basic
configurations, τe = 0). In fact, figure 17 shows that an LOES match over the frequency range of 0.1- 10
rad/sec would produce -τe < 0 (if negative time delays were allowed) for Configurations 1A, 2A, and 7A.
The small positive τe for Configurations 1C and 2B results from the relatively low frequency of the secondorder lag (ω3) for these cases, 16 rad/sec as opposed to 63 rad/sec.
There are two potential methods for dealing with lead/lag systems like those of table XIII; unfortunately,
neither is physically very appealing. And in each, there is an underlying question as to the universality of
the equivalent systems approach.
a. Redefine Limits on ζe
If ζemin for Level 1 were increased from 0.35 to 0.50, four configurations in table XIII would fit the
requirements (ignoring Configuration 1C, for which none of this discussion is applicable). But restricting
unaugmented vehicles as well is not appealing since lower ζsp is very well supported by flight test data for
classical aircraft. The alternative to specify two sets of requirements -- one for unaugmented aircraft and
another for augmented aircraft, is especially unattractive, since this is tacit admission that equivalent
systems is a misnomer. Additionally, it presents the problem of defining the specific level of augmentation
at which the requirements would change over. For example should addition of a simple high-frequency
stick filter (whose only major effect is to increase τe, figure 18) change the requirement? In fact, the
problem with Configurations 1A, 2A, 2B, and 7A is directly traceable from pilot commentary to
overabruptness and apparently has nothing to do with damping ratio at all. In figure 16, the high gain at
high frequency would indeed be expected to lead to an abrupt response. If for some unforeseen reason a
flight control system designer would ever suggest equalization that would produce such humps in the
frequency response, use of a frequency-response criterion is suggested. Four of the five configurations of
table XIII fit the bandwidth requirement (see figure 64).
186
MIL-STD-1797A
APPENDIX A
FIGURE 16. Comparison of Bode amplitude
plots for basic and augmented configurations
of table XIII.
FIGURE 17. Comparison of phase angles for
basic and augmented configurations of
table XIII.
b. Redefine τe
As mentioned above, three of the four low-ζe violators of table XIII also have τe = 0. As figure 17
suggests, a better LOES fit is obtained for these three cases if τe is allowed to be less than zero.
Specifically, negative time delays can be found in an LOES match to be as follows:
Configuration 1A --τe = -0.004 sec
Configuration 2A --τe = -0.008 sec
Configuration 7A --τe = -0.014 sec
Physically, unrealizable negative time delay, or time lead, might be considered to represent a HOS which
is too abrupt (i.e., if τ < 0, the system responds to an input τ seconds before the input is made or has finite
magnitude at zero time) more or less in keeping with the above-noted pilot commentary.
187
MIL-STD-1797A
APPENDIX A
FIGURE 18. Effect of first- and second-order lags on equivalent time delay and pilot rating:
LAHOS configurations (AFFDL-TR-78-122).
SIGNIFICANCE OF CAP OR ωsp AND n/α
For conventional aircraft, without direct lift control, the piloting task can be viewed as an inner- and outerloop, single-controller task as indicated for equivalent-system dynamics in figure 19. This figure points out
that 1/Tθ2 plays a role in the short-term attitude dynamics and also defines the short-term flight path (or
load factor) response. The boundaries in figure 13 are based on a combination of load factor response
and the quickness of the pitch attitude response to a control input, i.e., ωsp /(n/α). Some SP
2
interpretations of this parameter are given in the following discussion (In figure 19 note that γ/θ, being the
ratio of bare-airframe numerators, incorporates the actual value of 1/Tθ2, not necessarily the same as the
equivalent system's 1/Te if equalization is employed in feedback or feedforward loops.
a. Control Anticipation Parameter (CAP)
Bihrie in AFFDL-TR-65-198 defines the Control Anticipation Parameter (CAP) as the ratio of initial
pitching acceleration to steady-state normal acceleration (the pseudo steady state corresponding to the
two-degree-of-freedom short-period approximation).
&θ&
CAP ∆
= ∆n
zss
rad / sec 2
g
=&
ω 2sp
V 1
g Tθ2
188
(rad / sec) 2
( ft / sec)
(1 sec)
( ft / sec 2 )
or g-1sec-2
MIL-STD-1797A
APPENDIX A
γε
γc +
-
Yγ
θc +
θε
-
Yθ
F es
θ
Fs
F ee l
S y s tem δ es
and
A c tu a to r s
+
A θ (Tθ2 )e
− τe
θ
1 / Tθ 2
γ
=
θ
( s + 1 / Tθ 2 )
γ
θs
(0 ) ζ e , ω e
e
γ
Fs
M δ ( s + 1 / Tθ 2 )
θ
es
=
δ es
s (s 2 + 2 ζ s p ω s p s + ω s2p )
+
e
−τ
A γ e eγs
(0 ) ζ e , ω e
where
(a) = (s + a)
[ζ, ω] = [s 2 + 2ζωs + ω 2 ]
FIGURE 19. Pilot control of pitch attitude and flight path.
the last expression an approximation holding generally for aircraft with negligible control system dynamics
and tail lift effect, as is common, where (V/g) (1/Tθ2) likewise can be approximated by n/α. Because of the
time lapse before reaching the steady state, a pilot needs an earlier indication of the response to control
inputs - and both the initial and final responses must be neither too sensitive nor too insensitive to the
commanded flight-path change.
Note that &θ& o and ∆nzss apply to the time response to a step input, for an actual system of any order. If &θ& o
is interpreted as the maximum (as discussed later) and there is a nonzero ∆nzss, then that interpretation
has no call to determine any individual equivalent system parameters such as ζsp.
b. Frequency response interpretation
Equivalently, in the frequency domain the high-frequency gain of pitch acceleration (thought to be
important in fine tracking tasks) is given by MFs and the steady-state gain of normal load factor (thought to
be important in gross, or outer-loop, tasks) by MFs (n/α)/ ωsp . Hence their ratio is CAP (see figure 20).
2
c. MIL-F-8785B BIUG interpretation
A closely related analysis in AFFDL-TR-64-72 arrives at (FS/n) MFs equal to ωsp /(n/α) where MFs is the
2
initial pitch acceleration per pound. This can be seen in the asymptotic nz/Fs frequency response of figure
20.
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FIGURE 20. Definition of CAP from frequency response asymptotes.
d. Maneuvering stability margin interpretation
Because n/α is proportional to CLα and, roughly, ωsp to Cmα, ωsp /(n/α) is widely recognized as being
2
2
related to static margin, -dCm/dCL. In fact,
n
α
=&
C Lα q S
W
and
ω 2sp
=&
 C mα

− q Sc
ρSc
C Lα 
+
⋅ C mq 


Iy
4m
 C Lα

Therefore,
ω 2sp
nα
=&
−c ⋅
W
Iy
 C mα

ρSc

+
⋅ C mq 
 CL

4(W S )
α


= hm ⋅ g
(k c )
2
y
where hm is the maneuver margin expressed as a fraction of c (i.e., hm is the distance, in chord lengths,
of the maneuvering neutral point aft of the c.g.), and ky is the nondimensional pitch radius of gyration.
(With Cmq neglected, hm reduces to one common, simple definition of static margin, hs).
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For many aircraft, ky is about 17 percent of the aircraft length l , so ωsp /(n/α) = 1100 hm c / l 2 rad2/sec2.
2
For the F-4 aircraft, with 64 ft length and 16 ft c, the specified Level 1 value of 0.28 for ωsp /(n/α) reduces
2
to a stick-fixed maneuver margin requirement of 6.5 percent. Thus this requirement is comparable to the
earlier 5 percent static margin requirement in U.S. Air Force Specification 1815B/Navy Specification SR119B.
e. Importance of ωsp and (n/α) individually
Rationally, there should exist lower limits on satisfactory short-period frequency and on normalacceleration sensitivity to pitch control. This concept is consistent with the in-flight and ground-based
simulator experiments of AFWAL-TR-81-3118 and Mooij & van Gool and Wilhelm & Lange in AGARDCP-333 (Gibson opts as some others have done for bounding 1/Tθ2 instead of n/α). Although we have
retained the numerical values of MIL-F-8785C, for which there still are few data points, Mooij & van Gool
indicate more stringent limits.
POSSIBLE MODIFICATIONS TO CAP
As noted by Bischoff (NADC-81186-60), the control anticipation parameter must be redefined for aircraft
with effective time delay since &θ& o = 0 in this case. Following DiFranco (AFFDL-TR-66-163), Bischoff
defines, on the basis of a unit step stick force input, a more general control anticipation parameter, CAP ′ ,
as
&θ&
max HOS
CAP ′ ∆
=
n z ss
The maximum pitch acceleration, &θ&maxHOS , will occur sometime after the force input as shown on figure
21. The parameter CAP ′ is further extended to the short-period lower-order equivalent system model by,
defining
CAPe′
=&
2
&θ&
 ω2sp 

 ⋅ max HOS ∆ ωeff
= (n α )
&&
n α
e

 e θLOES t = τe
where e denotes LOES parameters. In this form, CAPe′ is a hybrid frequency response and timeresponse parameter easily determined from HOS responses. Bischoff claims "similar results... for the
higher order fixed and free low order equivalent system." The [ ω 2sp /(n/α)]e alone does not give a good
approximation to CAP ′ , because the short-period LOES model will not generally be accurate in the high
frequency region which largely determines the initial pitch acceleration history for a step input. Thus
2
as determined from the HOS response (as on figure 21).
[ ω /(n/α)]e is modified according to &θ&
sp
max
Generally, this CAP ′ will vary with the magnitude of the input, because of actuator rate limiting.
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FIGURE 21. Pitch acceleration response to a unit step force input.
NADC-81186-60 accounts for time delay explicitly by defining flying qualities Levels in the CAP ′ - τ
plane, as shown on figures 22 and 23. The boundaries shown for each flying quality Level were defined
by correlations of data from DiFranco (AFFDL-TR-66-63). Neal and Smith (AFFDL-TR-70-74), and the
LAHOS study (AFFDL-TR-78-122). These boundaries do seem to correlate the data that Bischoff plotted
slightly better than does the present requirement based on CAP (compare to figures 36 and 38), and
Bischoff shows excellent correlation between CAP ′ and CAPe′ for these data. However, the CAPe′
parameter is subject to all the limitations for equivalent systems noted in this requirement. Hence most of
the points that do not correlate with CAP or ωsp Tθ2 will also be missed by CAPe′ and CAP ′ . The
bandwidth specification appears to do a somewhat better job than CAP ′ .
LEVEL 3
All the suggested short-term pitch response requirements share a common Level 3 floor in recognition of
(a) the demonstrated controllability of somewhat unstable airplanes and helicopters from their beginning
days and (b) our inability to come up with anything better at this time.
A first-order divergence (T2 = 6 sec) is allowed for the Level 3 pitch attitude dynamics. This is consistent
with the Level 3 static stability requirement in 4.4.1. The 6-second limit on instability was derived from inflight and ground-based simulator studies which have documented the degree of instability that is safely
flyable. AFFDL-TR-72-143, for example indicated a Level 2 boundary with T2 (based on the unstable
aperiodic root) of 2.5 seconds in light turbulence and 4.25 seconds in moderate turbulence. Pilot ratings
were fairly constant at 5 to 6 until the time to double amplitude was reduced below 6 seconds, when
significant deterioration began.
Some margin is allowed in order to account for pilot distraction, design uncertainties, etc. In this region
near neutral stability, the root locations generally are extremely sensitive to static margin. On the other
hand, Schuler (AFWAL-TR-82-3014) concludes from his ground-based simulations that for light total
damping (small value of the other, stable short-term pole) the 6 second limit can be unconservative.
SUPPORTING DATA
Supporting data for both CAP and ωsp Tθ2 forms of equivalent-system criteria, and for the Level 3
requirements, are presented in the discussion of the ωsp Tθ2 criteria which follows.
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APPENDIX A
FIGURE 22. Time delay versus CAP′ - Neal-Smith data
FIGURE 23. Time delay versus CAP′ - LAHOS data (from
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APPENDIX A
(from NADC-81186-60).
NADC-81186-60).
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APPENDIX A
B. ωsp Tθ2 , ζsp , τθ
The equivalent pitch rate and normal load factor transfer-function forms for pilot control-deflection or
control-force inputs
δ es
q(s)
or Fes (s)
δ es
[s
=
(
K θ s s + 1 Tθ1
2
+ 2ζ p ωp s + ωp2
(
) (s + 1 Tθ2 ) e −τes
] [s
2
+ 2ζ sp ω sp s + ω 2sp
)
K n s s + 1 Th1 e − τns
n ′z (s)
=
or Fes (s)
s 2 + 2ζ p ωp s + ωp2 s 2 + 2ζ sp ω sp s + ω 2sp
[
][
]
]
simultaneously shall be fit to the corresponding actual response of the aircraft over a frequency range of
0.1 to 10 radians per second. The parameter n′z is normal acceleration at the instantaneous center of
rotation for pilot pitch control inputs. Corresponding matches shall be made for pilot control-deflection
inputs. The requirements of figure 24 apply to the equivalent-system parameters determined from the
best match for force inputs, and also for deflection inputs; the procuring activity will be the judge of the
adequacy of the matches.
For Level 3, T2, the time to double amplitude based on the value of the unstable root shall be no less than
6 seconds. In the presence of any other Level 3 flying qualities, ζsp shall be at least 0.05 unless flight
safety is otherwise demonstrated to the satisfaction of the procuring activity. T2 applies to the value of an
unstable root: T2 = -(ln 2)/λ where λ, is the value of the unstable root.
Requirements on the equivalent pitch time delay, τθ, apply to the value for θ(s)/δes(s) for a deflection
control system (pilot controller deflection commands the control effectors) and to θ(s)/Fes(s) for a force
control system (pilot controller force commands the control effectors):
Level
Allowable Delay
1
0.10 sec
2
0.20
3
0.25
If an adequate match cannot be found, the contractor with the concurrence of the procuring activity shall
substitute an alternative requirement.
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APPENDIX A
FIGURE 24. Requirements for short-term pitch response to pitch controller (ωspTθ2 vs ζsp).
196
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APPENDIX A
Equivalent systems are discussed above, in connection with the CAP. Much of that discussion was based
upon the premise that the normal acceleration response to attitude changes is a primary factor affecting
the pilot's perception of the minimum allowable ωsp. It is, of course, also true that the pitch attitude
response to pitch control inputs is in itself of great importance. Whether the appropriate correlating
parameter is n/α or 1/Tθsp unresolved: data that are correlatable with 1/Tθ2 will generally also correlate
with n/α. This issue was studied in AIAA Paper 69-898, where it was observed debatably that the product
ωsp Tθ2 provided a slightly better correlation than CAP. Physically, for the classical case ωsp Tθ2 represents
the lag in phase (at ωsp) or time between aircraft responses in pitch attitude and path. If 1/Tθ2 is too large
with respect to ωsp or the closed-loop pitch bandwidth, the path and attitude responses may not be
separated enough to give a pilot the additional cues he needs in order to control the outer, slower path
loop. The aircraft responses in attitude and flight path to elevator deflection occur almost simultaneously,
resulting in abrupt heave responses to the pitch controller. This produces pilot comments such as "trim
hard to find" and "pilot effort produces oscillations.” However, too great a frequency separation creates a
& /δ|, manifest as a large pitch rate overshoot, or bobbling tendency in closed-loop
large hump in | θ
tracking. We see that ωsp Tθ2, in combination with ζsp, also defines the shape of the attitude frequency
response: log (ωsp Tθ2) is the difference in frequency, on the usual logarithmic scale, between ωsp and Tθ2.
Desirable values yield a K/s shape of θ/δ in the frequency range of primary interest (see AIAA Paper 69898). A useful criterion, therefore, is the product ωsp Tθ2.
SUPPORTING DATA
The data base consists of airplanes with classical flying qualities as well as highly augmented airplanes
which are treated in this section by reduction to lower-order equivalent systems. The supporting data for
classical airplanes and highly augmented airplanes are presented separately in the following two
subsections, which are further subdivided according to the Flight Phase Category of the data. In each
instance, both the CAP and the ωsp Tθ2 forms of equivalent systems criteria are discussed. Lastly, the
Level 3 substantiation is presented.
Supporting data--classical airplanes
Most of the available data are for Category A Flight Phases only. A small amount of Category C data is
available, while data for Category B are extremely sparse. There is a considerable amount of existing
data which do not support the boundaries in figures 13 and 14 (see AFFDL-TR-69-72). However, a close
review of the data reveals that most of the scatter was due to secondary effects. For example, in some
cases the stick force per g (Fs /n) was outside the Level 1 limits. In other cases the tests were performed
with an extremely low load factor limit (nz < 2.0 g), or with the shore-period frequency near a wing
structural mode. There is evidence in the references that in these cases the extraneous factors may be
influencing pilot ratings.
A careful review of AFFDL-TR-66-63, FDL-TDR-64-60, WADC-TR-55-299, WADC-TR-57-719, AFFDLTR-68-91, NASA-TN-D-779, NASA-TM-X-1584, NASA-TN-D-3971, WADC- TR-54-594, AFFDL-TR-69-3,
Boeing Report D6-10725, Cornell Report TB-1444-F-1, and Princeton University Report 777 was
undertaken to cull out inappropriate data. Those reports which were complete enough to allow a thorough
analysis of the test conditions and results were reviewed in detail. Others were considered to raise too
many questions to be analyzed with confidence. (This is not meant to imply that some of the reports are
invalid, but that they were not complete enough to gain sufficient insight into the causes for expected or
unexpected pilot ratings.)
In particular, valid and usable reports were those which contained at least the following: 1) characteristics
of short-period mode(s); 2) description of aircraft actuators, feel system, etc.; 3) description of maneuvers;
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MIL-STD-1797A
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4) flight conditions; 5) pilot opinion rating scales used; and 6) pilot comments or discussion of pilot
comments. The last factor especially reduced the number of reports retained for analysis. Pilots'
descriptions of motions, responses, flight conditions, and control forces were considered essential to
justify any pilot ratings which were inconsistent with other test data or with expectations.
WADC-TR-55-299. WADC-TR-57-719, NASA-TN-D-779, WADC-TR-54-594, Boeing Report D6-10725,
and Princeton University Report 777 do not contain sufficient pilot commentary, if any, to be useful in the
above context. In addition, high Mach number data in NASA-TM-X-1584 were not used because pilots
considered the attitude display of the aircraft (an XB-70) to be inadequate when operating at Mach 3.
Low-n/α tests of AFFDL-TR-66-163 (n/α = 16.9 g/rad) were subject to a buffet-onset load factor limit of nz
= 2 g--low for evaluating a fighter-type aircraft; was also noted (AFFDL-TR-66-163, page 41) that:
Airplane sensitivity was more erratic and difficult to control when the structural modes of the
airplane were excited. The primary mode excited was wing bending, which occurred at
frequencies between 17 and 21 rad/sec (2.7 to 3.3 cps). These bending frequencies were
observed in the oscillograph record of a wing tip mounted accelerometer and are a function of
the fuel remaining in the tip tanks. Both pilots commented on the varying degree of structural
excitation that occurred when the airplane undamped frequencies varied from approximately
8 to 11.5 rad/sec (approximately half the structural frequencies). The erratic nature of the pilot
ratings and pilot-selected stick forces in this region are also understandable. The pilots were
obviously correcting and interpreting sensitivity due to structural factors as well as the
inherent airplane sensitivity.
However, Chalk has stated (personal communication) that the evaluation pilots generally seemed to be
able to discount that effect to give valid ratings.
Based on the evidence, some data of AFFDL-TR-66-163, AFFDL-TR-70-74, and AFFDL-TR-69-3 (which
are T-33-based experiments) with ωsp > 8 rad/sec are presented here. We note that all three reports show
a rating deterioration at high ωsp that tends to support the given upper bound Ratings data from Cornell
Report TB-1444-F-1 (taken in a B-26, simulating a fighter configuration) showed greater scatter and
overall better (lower) ratings than any of the other reports. This led to an evaluation of the reference, and
to the conclusion that the tasks of TB-1444-F-1 were not sufficiently demanding to provide a good basis
for evaluation of closed-loop handling qualities. Hence the data were not used.
In summary, AFFDL-TR-66-163, AFFDL-TR-70-74, AFFDL-TR-68-91, and AFFDL-TR-69-3 provided
good short-period data for Category A; NASA-TM-X-1584, NASA-CR-159236, and AFWAL-TR-83-3015
contain usable Category B data; FDL-TDR-64-60 and NASA-TN-D-3971 contained Category C
information; for large aircraft AFFDL-TR-72-41, NASA-CR-159236, AFWAL- TR-81-3118, AFWAL-TR-833015, and AGARD-CP-333 furnish additional insight.
Category A
1. ωsp /(n/α) criterion
2
Figure 25 shows the short-period frequency boundaries for the Category A Flight Phases. The applicable
data (with Level 1 FS/n) from AFFDL-TR-66-63, AFFDL-TR-68-91, and AFFDL-TR-69-3 are compared
with the boundaries. These data represent 52 separate ωsp - n/α combinations flown and rated b), six
pilots. Eight configurations which fell within the Level 1 boundaries were rated Level 2 or worse. The
boundaries correctly predicted pilot ratings about 80 percent of the time--an adequate success rate given
the variability of flight tests and pilot ratings. Note that most of the violations occur at large n/α (as at high
speed). The Level 1 boundaries are slightly more lax than the best fit to the data presented.
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MIL-STD-1797A
APPENDIX A
The data in figure 25 represent those cases for which ωsp and Fs/n were within the present Level 1
boundaries. Therefore, the ratings shown can be assumed to be due solely to short-period frequency and
n/α influences. For some experiments it could be argued that even Level 2 Fs/n should be plotted, since
the pilots were allowed to select the optimum value. We have taken a somewhat conservative approach
by eliminating these data. Our reasoning was that Levels 2 and 3 are boundaries for an off-nominal or
failed state, and that pilots will not have a chance to optimize Fs/n after a failure. It should be noted that
when Level 2 values of FS/n are selected by pilots it is usually to account for a basic flying quality
deficiency. For example, a pilot would desire a very low FS/n after a failure which results in a statically
unstable airframe, requiring pulse-like control inputs. The data generally support the requirements,
although the Level 1 lower boundary may be somewhat low.
The baseline cases of AFFDL-TR-70-74, discussed under Augmented aircraft, also support the Category
A boundaries.
FIGURE 25. Comparison of pilot ratings with category A short-period frequency requirements.
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MIL-STD-1797A
APPENDIX A
Short-period damping, boundaries are shown on figure 16. Since this criterion presents ζsp as an
independent parameter, it has been plotted directly against pilot rating. The other variables (ωsp , n/α and
Fs/n) are Level 1 values. The resulting data (48 individual ratings) show definite trends in correlation with
the boundaries, although there is a shortage of data points in the Level 2 and 3 areas. However, several
points for which damping is good (ζsp = 0.67-0.74) are rated Level 2 to 3. The low-damping data from
AFFDL-TR-66-63 suggest that the ζsp lower limits could be reduced. Seven ratings are worse (higher in
value) than predicted, all occurring within the Level 1 boundaries; a few are better than predicted. Any
possible interdependence between ζsp and ω 2sp /(n/α) can be taken into account by replotting the criterion
boundaries on a girid of ζsp vs. ω 2sp /(n/α) as in figure 27.
FIGURE 26. Comparison of pilot ratings with Category A short-period damping requirements.
The authors of AFFDL-TR-69-72 also noted that ζsp lower limits were too restrictive. However, the datasupported limits do not account for turbulence (it was minimal); therefore the MIL-F-8785B limits were
chosen somewhat higher.
The upper ω 2sp /(n/α) limits of MIL-F-8785C are difficult to confirm based on the figure 27 plot because the
validity of the high-frequency (ωsp > 8 rad/sec) data has been questioned, as noted above. Some of the
data may be usable, however, since the structural bending mode was reported to be most pronounced at
low speed with a high fuel load. For this Handbook we have used some of the basic high-frequency data
(unaugmented configurations). The upper boundary on ω 2sp /(n/α) has been retained.
For the C-5A airplane, AFWAL-TR-83-3015 presents (a) data showing generally Level 2 & 3 ω 2sp /(n/α),
figure 28, for two Category A tasks, terrain following and refueling – receiver, worse than Level 3
according to MIL-F-8785B/C at aft cg and (b) a letter from the 60th Military Airlift Wing (MAC) attesting
that “Most C-5A pilots find that the manual flying characteristics of the C-5A are excellent." Because some
Class III aircraft may be much smaller than the C-5A we do not wish to reduce the Category A
requirement generally for Class III aircraft. For very large aircraft, however, similar size or task artifacts
may affect tolerable levels of dynamic stability.
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MIL-STD-1797A
APPENDIX A
FIGURE 27. Comparison of pilot ratings with Category A short-period frequency
and damping ratio requirements.
2. ωspTθ2 vs. ζsp criterion
Figure 29 shows the data used to support the requirements based on ωspTθ2 and ζsp. The damping limits
are not supported by the pilot ratings but are consistent with the reasoning shown in the preceding
discussion. The absolute lower limits on ωsp utilized in the ω 2sp /(n/α) criterion have been retained in the
ωspTθ2 vs. ζsp requirement for the lack of any better data. They are presented in a table in figure 24a.
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MIL-STD-1797A
APPENDIX A
FIGURE 28. Category A C-5A flight test data.
FIGURE 29. Alternate Category A flying qualities
requirements for short-period pitch response.
202
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APPENDIX A
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MIL-STD-1797A
APPENDIX A
Two criteria proposed by Gibson in AGARD-CP-333 (section F), 2ζspωsp > 1/Tθ2 and Tθ2 > tγ, the
asymptotic time lag in flight-path response, are seen in figure 29 to correlate with the data as well as
ωspTθ2 does.
More work needs to be done to define the upper limits on ωspTθ2 for Category A. In AGARD-CP-333,
Gibson indicates good and poor regions of his variables “dropback/q” = Tθ2 –2ζsp/ωsp and qmax/q, a
function of ωspTθ2 and ζsp. ICAS-86-5.3.4 discusses both upper and lower limits of ωspTθ2, outlined in
Section F, related also to the response parameter of time to the pitch rate peak. All are task dependent
and most vary with Tθ2.
Category B
1. ω 2sp /(n/α) criterion
Applicable pilot ratings from NASA-TM-X-1584 (XB-70) are compared with the ωsp limits in figure 30 and
with the ζsp boundaries in figure 31. The data do not conflict with the boundaries.
We adopt NASA-CR-159236's recommendation to relax the ωsp /(n/α) floors based on Concorde cruise
2
and C-5A data. AFWAL-TR-83-3015's L-1011 data also tend to support such a change, considering the L1011 to have reasonably good flying qualities. The data are presented in figures 32 through 35.
2. ωsp Tθ2 vs. ζsp criterion
Since there are insufficient data to propose boundaries, the Category B limits have been made
compatible with the Category A and C limits, so ωsp Tθ2 = 1.0 for Level 1 and ωsp Tθ2 = 0.58 for Level 2.
Figure 36 illustrates the criterion, and compares the NASA-TM-X-1584 data.
Category C
1. ω 2sp /(n/α) criterion
The Category C flight tests data of FDL-TDR-64-60 and NASA-TN-D-3971 (T-33 and B-367-80,
respectively) and NADC-80157-60 (Navion) are compared with the short-period frequency requirements
in figure 37 and damping requirements in figure 38.
The frequency data fit the boundaries very well, with a success rate of about 81 percent. This is
comparable to the 80 percent for the Category A data, but there are far fewer Category C ratings, over a
much smaller range of n/α. However, ratings better than 3.5 seem to require ω 2sp /(n/α) above the
Category A Level 1 boundary, which is considerably higher than the corresponding Category C boundary.
Damping predictions are worse, 73 percent--identical to that for Category A. The data support reduction in
the minimum ζsp for all Levels, similar to those suggested by the Category A data. However, as for the
Category A data, these tests were conducted in minimal turbulence, so the MIL-F-8785C damping ratio
limits have been retained.
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MIL-STD-1797A
APPENDIX A
FIGURE 30. Comparison of pilot ratings with Category B short-period frequency requirements.
FIGURE 31. Comparison of pilot ratings with Category B short-period damping requirements.
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MIL-STD-1797A
APPENDIX A
FIGURE 32. Category B short-period
characteristics.
FIGURE 33. L-1011 climb short-period
characteristics.
FIGURE 34. L-1011 cruise short-period
characteristics (AFWAL-TR-
FIGURE 35. L-1011 descent short-period
characteristics (AFWAL-TR-83-
206
MIL-STD-1797A
APPENDIX A
83-3015).
3015).
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MIL-STD-1797A
APPENDIX A
FIGURE 36. Alternative Category B short-period flying qualities requirements
(NASA-TM-X-1584 data, Level 1 Fs/n).
Ratings of AFFDL-TR-78-122's baseline configurations, included under Augmented aircraft, support these
Category C ω 2sp /(n/α) boundaries.
Recommendations to relax the Category C ωsp boundary for large airplanes, as for example in AFWALTR-83-3015, have not been adopted for reasons expressed well by van Gool and Mooij in private
correspondence:
We realize that one of the reasons for changing Class III requirements is that the
currently flying large aircraft supposedly have Level 1 handling qualities while not
complying with many of the current requirements. Widening the boundaries may be only
part of the solution though. In our opinion the fact that the handling qualities of these
aircraft are satisfactory is mainly the result of an adapted piloting technique. The pilots
have learned to cope with low short period frequency, low acceleration sensitivity and
large time delays by avoiding to get into the control loop. Pilots flying aircraft like B 747,
DC 10 and C-5 will tell you that, e.g., the landing flare is an open loop maneuver. We
assume that the fact that time delays as high as 0.4 s did not affect the pilot ratings in the
Lockheed study was caused by the use of an open-loop landing technique.
For military operations this open-loop technique appears inadequate.
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MIL-STD-1797A
APPENDIX A
FIGURE 37. Comparison of pilot ratings with Category C short-period frequency requirements.
Supporting data--large aircraft (Classes II and III)
There have been frequent suggestions by manufacturers of large aircraft that the Level 1 lower
requirements on ωsp vs. n/α should be lowered in the landing approach Calspan, in NASA-CR-159236,
did not recommend this relaxation because data in the BIUG, AFFDL-TR-69-72, and in Calspan's
proposal to modify MIL-F-8785B, AFFDL-TR-72-41, substantiate the original boundary quite well. Informal
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MIL-STD-1797A
APPENDIX A
discussions with pilots of large aircraft indicate that current large airplanes possess generally comfortable
bandwidth for routine use. However, when the task difficulty increases due to weather conditions, for
example, hard landings and go-arounds are common. In view of the very demanding landing conditions
being imposed on large military aircraft such as the YC-14, YC-15 and C-17, relaxation of the requirement
seems imprudent until more substantiating data become available. Indeed, although AFWAL-TR-83-3015
indicates that somewhat lower values can be satisfactory, AFWAL-TR-81-3118 and Mooij & van Gool in
AGARD-CP-333 indicate that the lower Level 1 boundary might be raised.
FIGURE 38. Comparison of pilot ratings with Category C short-period damping requirements.
However, a body of data supports relaxation of the Level 2 & 3 requirement. Consistent with allowing a
divergence in the Level 3 case, that floor has been removed here. A lower Level 2 floor of ω 2sp /(n/α) =
0.05 suggested for transport aircraft in NASA-CR-159236 has been adopted. The data in figures 39 and
40 tend to support this new bound.
For lower-bandwidth tasks, the recommended maximum time delays seem too strict. WADC-TR-56-258
relates AFFDL-TR-78-122, AFFDL-TR-70-74, and AFFDL-TR-68-90 data according to “The closed-loop
pitch attitude bandwidths which pilots were generally believed to be requiring in these experiments.” The
effective time-delay boundaries are shown to correspond to constant phase lags at those bandwidth
frequencies (figure 41) [Although this effective time delay comes from a time history (see “C. Transient
Peak Ratio,..."), equivalent time delay τe would follow the same trend.] Considering that for AFWAL-TR81-3118's million-pound transport "The landing approach and simulated touchdown task ... with a large,
slow-responding aircraft can be considered as having the same bandwidth requirements (1.5 rad/sec) as
the fighter up-and-away and low altitude waveoff task of AFFDL-TR-68-90," the trend is verified.
However, the data are too sparse to be definitive. Other data, cited in AFWAL-TR-83-3015, suggest even
less stringent limits. And for emergencies, an experimental Airbus has been flown unstable
(unaugmented) for five minutes through the pitch trim system, with a time delay of over a second--but
without attempting approach or landing. Considering higher-gain tasks and a possibly more severe
environment, we cannot recommend such large relaxations. Again quoting van Gool and Mooij,
Our experience with time delays lead us to believe that for closed-loop control of
transport aircraft in approach and landing the Level 1 boundary on equivalent time delay
(determined with the equivalent system technique) is higher than the Mil Spec value of
0.1 s, but it will certainly not be as high as 0.4 s (in NLR experiments we obtained
satisfactory pilot ratings with 0.25 s equivalent delay in pitch and roll control).
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211
MIL-STD-1797A
APPENDIX A
FIGURE 39. Pilot ratings for large airplanes (nominal equivalent short-period
parameters) from AFWAL-TR-81-3118.
212
MIL-STD-1797A
APPENDIX A
FIGURE 40. Category C flight data for the Lockheed C-5A and L-1011, AFWAL-TR-83-3015.
213
MIL-STD-1797A
APPENDIX A
214
MIL-STD-1797A
APPENDIX A
FIGURE 41. Time delay bands associated with flying qualities boundaries vs
bandwidth, AFWAL-TR-81-3118.
2. ωspTθ2, ζsp criterion
Preliminary, straight-line boundaries of ωspTθ2 and ζsp are shown on figure 42 (solid tines). The data fit
these limits with a confidence level of about 82 percent, so the boundaries seem to work well. An even
better fit is given by the dashed lines on figure 42, which correlate with more than 90 percent of the
ratings. Note that these latter boundaries tend to eliminate combinations of low damping and low
frequency. We have therefore elected to set the criterion boundaries based on the dashed lines on figure
42. The minimum levels of ωsp (independent of ωspTθ2, taken directly from the ωsp, n/α criterion, are
presented in a table on figure 25c. The minimum levels of 1/ Tθ2 are based on the n/α of limits in figure
13c by assuming an approach speed of 135 kt and noting that 1/Tθ2 = (g/V)(n/α).
215
MIL-STD-1797A
APPENDIX A
FIGURE 42. Alternate Category C short-period flying qualities requirements.
216
MIL-STD-1797A
APPENDIX A
Gibson's (AGARD-CP-333) time-response criteria are seen to correlate well. The more complete criteria
in ICAS-86-5.3.4 indicate that the agreement is associated with flight path delay of 2 seconds for the large
B-367-80 (NASA-TN-D-3971) and with sluggish attitude response for the small T-33 (FDL-TDR-64-60)
even though the path delay is only 1 second. For the examples given, a minimum Level 1 ωspTθ2 value of
1.8 is suggested, limiting the time to peak pitch rate to 2 seconds (large aircraft NASA-TN-D-3971) and to
1.1 seconds (Class 4 aircraft, FDL-TDR-64-60). This minimum would increase to 2.8 for many Class 4
aircraft with Tθ2 of nearly 2.0. The limited Navion data, not shown, do not correlate as well.
Supporting data--augmented aircraft
In this section, HOS and LOES will be examined and compared with the specification boundaries of figure
13. The sources of the HOS data are two Calspan research efforts, by Neal and Smith (AFFDL-TR-70-74)
for Category A Flight Phases and by Smith (AFFDL-TR-78-122) for Category C. The matching was
generally done only on the pitch response, with 1/Tθ2 fixed.
Neal-Smith data (AFFDL-TR-70-74)
The in-flight NT-33 experiments conducted by Neal and Smith represent a first look at generic variations
typical of highly augmented aircraft. For evaluation of HOS characteristics and criteria, 51 separate FCS/
short-period configurations were flown on the USAF/Calspan NT-33. Of these, some require
qualifications: tests conducted at 250 kt (n/α = 18.5) were limited to a load factor of 2.5 g, due to buffet
onset; Neal and Smith reported that the pilots did not fly these tests as aggressively as they did the highspeed (350 kt) tests.
Figure 43 shows the equivalent dynamics of the 51 Neal-Smith configurations (from MDC Rpt A6792) and
corresponding Cooper-Harper ratings. The frequency and damping ratio limits have been cross-plotted to
facilitate presentation of the data. Figure 43a includes actual pilot ratings for each pilot; for those cases
which have τe< 0.1 (Level 1), correlation is quite good. There is clearly a relationship between τe and PR
(figure 43b), though the τe limits appear to be too lenient since many of the configurations which are
predicted to have Level 1 ζsp and ωsp have higher (poorer) ratings than predicted by τe alone. In figure 43b
only the mean pilot rating and standard deviation have been plotted, to reduce the number of data points.
A point-by-point comparison of the LOES data on figure 43 shows that the flying qualities Levels are
accurately predicted for about two-thirds of the configurations. (This requires some liberal interpretation,
e.g., if a PR change of one-half rating would improve correlation, the configuration is assumed to fit the
criterion. Such a PR variation is well within the range of normal ratings variations.) This correlation rate is
not outstanding, but is close to that found for flying qualities data in general (for example, the data used in
AFFDL-TR-69-72 to define the ζsp and ωsp boundaries of MIL-F-8785B). Several Level 2 rated
configurations lie well within the Level 1 boundaries. However, the four configurations with ζe =
& 0.4
indicate that if the damping ratio lower limits were increased correlation would improve. This is in
contradiction with the results of the previous section, where a decreased ζsp limit would improve
correlation for classical aircraft.
Not surprisingly, the ωspTθ2 criterion does not correlate the AFFDL-TR-70-74 data any better (figure 44).
Correlation would be improved if the minimum damping for Level 1 were increased to 0.5 and if ωspTθ2
were increased to a value around 1.85. Both changes would be incompatible with the data base for
classical unaugmented aircraft.
217
MIL-STD-1797A
APPENDIX A
FIGURE 43. Comparison of LOES dynamics with short-period requirements; Category A,
Neal-Smith (AFFDL-TR-70-74) configurations, MCAIR (MDC Rpt A6792) matches.
218
MIL-STD-1797A
APPENDIX A
FIGURE 44. Comparison of Neal-Smith LOES characteristics with ωspTθ2 vs ζsp
The inconsistency between the unaugmented and augmented airplane data bases needs further study.
Until such analyses can be conducted, we have elected to utilize the MIL-F-8785C boundaries, which are
based on classical airplane data. However, for the purposes of guidance, according to the data the
equivalent frequency and damping for Level 1 augmented aircraft should meet the following criteria for
Category A Flight Phases:
ωspTθ2 ≥ 1.85 or 3.6 ≥ ω 2sp / (n/α) ≥ 0.37, ζsp ≥ 0.50
219
MIL-STD-1797A
APPENDIX A
LAHOS data (AFFDL-TR-78-122)
A systematic evaluation comparable to the Neal-Smith study was conducted for HOS effects in landing
approach (Category C). The Landing Approach Higher Order System (LAHOS) study (AFFDL-TR-78-122)
provides a good set of data for comparing LOES with Category C requirements.
Figure 45a compares LOES matches with ω 2sp /(n/α) vs ζsp , with τe indicated; and the data are plotted
against the allowable time delay (τe) requirements on figure 45b, with CAP and ζsp noted. The LAHOS
data correlate well with the boundaries. In fact, the flying qualities of about 85 percent of the LAHOS
configurations are accurately predicted. The only area of poor correlation on figure 45 involves those
configurations which should have Level 1 flying qualities, but are rated by the pilots as Level 2. This may
be in part a function of the fidelity of the tests and the realism of the tasks: a combination of instrument
and visual approaches through touchdown and landing, or with intentional go-around maneuvers. Most of
the classical data upon which the short-period requirements are based (see AFFDL-TR-69-72) were
generated for approach and go-around tasks only, seldom including actual landing, which is normally the
most critical area. The LAHOS data may therefore be more representative of flying qualities in the
terminal phases, of flight. AFFDL-TR-78-122 discusses this at some length.
One shortcoming of LAHOS is that the equivalent systems do not cover a wide range of ζsp and ωsp
(figure 45a); these are Level 2 or worse for only nine of the 46 configurations. LAHOS is primarily an
exercise of the τe limits (figure 45b). (This is not a shortcoming of the LOES approach, but an artifact of
the range of HOS evaluated in the LAHOS program).
Not surprisingly, ωsp Tθ2 vs. ζsp is very similar (figure 46). Some improvement in the correlations would be
possible by increasing the ωsp Tθ2 limit to 1.85. There is one data point to suggest a possible increase in
ωsp Tθ2 to 2.2.
Gibson (AGARD-CP-333) recommends cutting off the lower right-hand corner of figure 36's Level 1
bound (ω > 2ζ for his tγ < Tθ2) to eliminate negative dropback.
Level 3
Substantiation for allowing an aperiodic divergence with a time to double amplitude of 6 sec for Level 3
flying qualities was given in AFWAL-TR-81-3109. The essence of that discussion is repeated below.
In response to a pulse control input, stable aircraft reach steady values of α, h and V; unstable aircraft
have the same initial response, then diverge, as illustrated by figure 47 (from NASA-TM-X-62). For a
supersonic transport design, impulse responses are shown for various degrees of static instability as Cmα,
is varied. Also shown is the response of a configuration having much more static instability, with time to
double amplitude reduced by a pitch damper. Evaluation pilots rated both of these configurations
unacceptable, but termed the latter's characteristics insidious. On the other hand, Schuler's (AFWAL-TR82-3014) ground-based simulation showed additional total damping to be helpful. From NASA-TN-D-173,
commenting on an F9F-2 airplane with static instability ameliorated by a damper to give about 6 seconds
to double amplitude:
The rate of divergence of the airspeed was scarcely noticeable to the pilots in normal
flying. However, this degree of instability might be objectionable for flight operations
where accurate control of airspeed is required.
220
MIL-STD-1797A
APPENDIX A
FIGURE 45. Comparison of LOES dynamics with short-period requirements; Category C, LAHOS (AFFDL-TR-78-122)
configurations, MCAIR ("Equivalent Systems Approach for Flying Qualities Specification").
221
MIL-STD-1797A
APPENDIX A
FIGURE 46. Comparison of LAHOS LOES characteristics with ωsp Tθ2 vs. ζsp.
From NASA-TN-D-779, pilot tolerance of aperiodic instability is much greater than of oscillatory instability
(figure 48). In that variable-stability YF-86D evaluation, an aperiodic divergence was not considered safe
with less than 1 sec to double amplitude: “there was a dangerous situation in that a short distraction of the
pilot's attention could allow the unstable vehicle to diverge to the point that it was difficult to recover." For
statically stable configurations “the unacceptable boundary is close to the zero damping boundary over
most of the frequency range...in the very low-frequency and very high-frequency ranges a small amount
of positive damping is required to remain within the acceptable region.” Commenting on this different
tolerance, Taylor and Day (NASA-TN-D-746) state:
222
MIL-STD-1797A
APPENDIX A
FIGURE 47. Comparison of effects of various stability characteristics on airplane response to elevator pulse
(-5 deg for 0.2 sec at t = 0) (NASA-TM-X-62).
223
MIL-STD-1797A
APPENDIX A
FIGURE 48. Contours of constant pilot opinion in statically unstable region;
constant stick-to-stabilizer gearing (NASA-TN-D-779).
At the higher frequencies, the technique for controlling the motion was not learned as
quickly...Controlling the pure divergence in the region of a static instability was more
natural and less tiring than controlling the oscillatory airplane motions, inasmuch as the
pilot need only counteract the angle-of-attack divergence without leading the motion to
stabilize the aircraft.
The unchanged phugoid requirement, T2 > 55 seconds for Level 3, still limits the low-frequency tolerable
oscillatory instability (the α, q, and nz feedbacks used in these variable-stability airplanes would not
suppress the phugoid mode in the region of low short-period frequency and damping). Higher-frequency
oscillatory instabilities are unlikely (except possibly through control-system failures), requiring
considerable negative aerodynamic damping; the limit of 6 seconds to double amplitude would fit the
Level 3 boundary of NASA-TN-D-779 for 0 < ω < 6 rad/sec.
For aperiodic instability, NASA-TN-D-211 shows that the boundary of acceptability for emergency
condition (Cooper 6.5) was insensitive to the value of lift-curve slope, or 1/Τθ2, or n/α, for positive lift-curve
slopes. This boundary value was 2 seconds to double amplitude.
224
MIL-STD-1797A
APPENDIX A
AFFDL-TR-72-143 demonstrates that at least at low speeds, the short-period approximation can give a
grossly incorrect value of T2. The T2 obtained from the angle-of-attack trace matched the three-degree-offreedom theoretical value fairly well when Cm(α) was actually linear and T2 was not too large. NASA-TMX-62 and AFFDL-TR-72-143 both elaborate on the range of values for time to double amplitude obtained
by different means: calculation from three-degree-of-freedom equations and various simplifications,
measurement from α, θ or V responses. M(α) nonlinearities gave different results for nose-up and nosedown perturbations; of course the worst direction would govern, for all reasonable magnitudes. Most of
the evaluations gave some consideration to turbulence. The AFFDL-TR-72-143 baseline configuration
had a Level 2 value of dγ/dV, but zero values were included in the evaluation--with a little improvement in
rating, but less noticeable in turbulence. The evaluations considered both visual and instrument flight.
On the basis of all these considerations, 6 seconds to double amplitude seems a reasonable, safe limit.
However, Schuler's fixed-base simulation (AFWAL-TR-82-3014) shows that the tolerable value of the
unstable root, λ1 , is affected by the value of the other, stable root, λ2 . The latter root must be at least a
certain minimum, but λ2 larger than that permits some increase in the instability of λ1. Operators may be
well advised to give pilots of potentially unstable aircraft some flight simulator experience with such
instability. It should be noted that pitch attitude and airspeed excursions will double in amplitude at
approximately the same rate, since (without autothrottle) u& =& gθ. Hence the allowable divergence in
attitude is the same as airspeed response to attitude (see 4.4.1).
It is desirable, though impractical at this time, to make the allowable instability a function of time. Clearly
an instability in cruise, where it might be hours before a runway is available, could be very tiring to the
pilot.
225
MIL-STD-1797A
APPENDIX A
C. Transient peak ratio, rise time, effective delay
The pitch rate response to a step input of pitch controller force, and also to a step controller deflection,
calculated from two-degree-of-freedom equations of motion (i.e., with speed constrained) shall exhibit the
characteristics defined in the following manner.
Two straight lines are drawn on the pitch rate time history (see figure 49) and the following measurements
are defined:
a. A horizontal line defining the steady-state pitch rate.
b. A sloping straight line tangent to the pitch rate time history at the point of maximum slope. This
line is extended to intersect the steady-state line and the time axis.
c. Time t1 measured from the instant of the step controller input to the time at the intersection of the
maximum-slope line with the time axis.
d. Time t2 measured from the instant of the step input to the time corresponding to the intersection
of the maximum slope line with the steady-state line.
e. The amplitude quantities ∆q1 and ∆q2 measured as follows: ∆q1 ∆
= maximum pitch rate minus the
steady state; ∆q2 ∆
= the steady state minus the first minimum.
The above-defined measurements shall meet the following design criteria, for step inputs of cockpitcontroller force and also for step inputs of cockpit-controller deflection, of magnitudes to produce both
small changes typical of fine tracking and large maneuvers up to limit load factor.
The equivalent time delay t1 shall be within the limits
Level
Equivalent time delay
1
t1 < .12 sec
2
t1 < .17 sec
3
t1 < .21 sec
For t1, use the step controller deflection for a deflection control system (pilot controller deflection
commands the control effectors) and the step controller force for a force control system (pilot controller
force commands the control effectors).
The transient peak ratio ∆q2 /∆q1 shall be equal to or less than the following:
Level
Max ∆q2 /∆q1
1
< .30
2
< .60
3
< .85
226
MIL-STD-1797A
APPENDIX A
The rise time parameter ∆t = t2 – t1 shall have a value between the following limits:
Nonterminal Flight Phases
Terminal Flight Phases
Min ∆t
Max ∆t
Min ∆t
Max ∆t
1
9/VT
500/VT
9/VT
200/VT
2
3.2/VT
1600/VT
3.2/VT
645/VT
Level
where VT is true airspeed, ft/sec.
FIGURE 49. Pitch rate response to step input of pitch controller force or deflection.
The product of the control-force gradient in steady maneuvering flight, Fs/n, and the maximum frequencyresponse amplitude ratio of pitch acceleration to pitch control force. | &θ& /Fs|max, shall not exceed the
following limits:
Level 1
3.6 rad/sec2/g
Level 2 and 3
10.0 rad/sec2/g
For Level 3, T2, the time to double amplitude based on the value of the unstable root, shall be no less
than 6 seconds. In the presence of any other Level 3 flying qualities. ζsp shall be at least 0.05 unless flight
safety is otherwise demonstrated to the satisfaction of the procuring activity. T2 applies to the value of the
unstable root: T2 = - (ln 2)/λ where λ is the value of the unstable root.
BACKGROUND
The time-response design criteria limit characteristics of the pitch rate response to pilot commands. This
format avoids explicit identification of dominant roots or equivalent system models by working directly with
227
MIL-STD-1797A
APPENDIX A
the pitch rate transient response. Nevertheless, by virtue of assuming a conventional-appearing response
these criteria do rely on a form of equivalent system. The criteria are applicable to aircraft exhibiting
conventional aircraft dynamic modes and to most pitch augmentation systems, but not to high-gain pitch
attitude command systems. If the residue of the phugoid or other low-frequency modes prohibits defining
a constant-speed, short-term, steady-state, pitch rate, it may be necessary to apply the criteria to the
pitch rate transient computed from constant-speed equations of motion. This discussion is adapted from
NASA-CR-159236.
The flight experiments reported in AFFDL-TR-68-90, AFFDL-TR-70-74, AFWAL-TR-81-3116, and
AFFDL-TR-78-122 have established the critical, detrimental nature of transport time delay and effective
time delay resulting from cascaded dynamic elements in the control system. The limits on effective time
delay are based on interpretation of the data in these documents.
Many calculation procedures have been proposed (see AFFDL-TR-70-155, AFFDL-TR-68-90, and
AFFDL-TR-70-74 for example). The methods developed in AFFDL-TR-68-90 require knowing the
coefficients of the characteristic equation of the higher-order system; the method of AFFDL-TR-70-155
requires a multivariable search to minimize the weighted sum of squares of errors of amplitude and phase
between the higher-order system and an assumed lower-order system having a time delay function e-as.
AFFDL-TR-70-74 reported analog matching of time-history responses using a lower-order transfer
function with a time delay; the best match was subjectively judged by the operator. These various
methods yield similar but different values of the effective time delay, and they are complex in
mathematical concept and application.
The effective time delay used in this time-domain requirement is defined on figure 49. It can be uniquely
defined (unless the response is dominated by higher-frequency modes or an aperiodic instability) and
easily evaluated either graphically or analytically. Values of t1 are generally smaller, for example, than the
time delay determined by the method of AFFDL-TR-70-155. It was necessary, therefore, to evaluate t1 for
the configurations evaluated in various experiments and to correlate these values with pilot rating.
These time-domain criteria are stated in terms of the transient peak ratio ∆q2 /∆q1. The intent is to ensure
adequate damping of the short-period or dominant mode of the pitch response. The specified values are
based on interpretation of short-period data in AFFDL-TR-72-41 and AFFDL-TR-69-72. For a classical
airplane response, the transient peak ratios would correspond to the following damping ratio values:
Level 1
TPR = .30
ζ = .36
Level 2
TPR = .60
ζ = .16
Level 3
TPR = .85
ζ = .052
The criteria also limit the effective rise time, ∆t, of the pitch rate response to a step pilot command (see
figure 49). The effective rise time is related to ω 2sp /(n/α) by the following:
ω2sp
nd
=
&θ&
0
nz ss
≅
qss ∆t
g
−
qss VT g
VT ∆t
Limits on ω 2sp /(n/α) are defined in MIL-F-8785C as a function of Flight Phase Category. For example, the
Level 1 limits for Flight Phase Category C are
.16 ≤ ω 2sp /(n/α) ≤ 3.6 rad3 /(g.sec2)
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MIL-STD-1797A
APPENDIX A
These limits can be related to the effective rise time as follows by substitution:
.16 ≤ g/(VT ∆ t) ≤ 3.6
By taking reciprocals, reversing inequality signs and rearranging, the following limits for ∆t result, in
commensurate units:
g/(3.6VT) ≤ ∆t < g/(.16VT)
This development indicates that limits on the effective rise time, ∆t, expressed as constants divided by the
true speed are analogous to the constant limits on ω 2sp /(n/α) used in MIL-F-8785C to specify short-period
frequency as a function of n/α. Separate ∆t limits are stated for terminal and nonterminal Flight Phases for
Level 1 and Level 2. The Level 3 limit is again 6 seconds to double amplitude.
SUPPORTING DATA
NASA-CR-159236 derives the numerical values of the ∆t limits for terminal Flight Phases directly from
ω 2sp /(n/α) limits of MIL-F-8785C. The maximum ∆t limits have been increased to accommodate flight test
data for existing aircraft such as the Concorde, XB-70, C-5A and from research data in “Recent Flight
Test Results on Minimum Longitudinal Handling Qualities for Transport Aircraft”. See figures 32 through
35. The Concorde data for cruise and the C-5A data for Flight Phase Category B in figure 32, the XB-70
test data on figure 50 together with the research data on that figure were used as the basis for reducing
the Level 1 and 2 minimum frequency limits for Flight Phase Category B of MIL-F-8785C. These
boundaries were then translated to maximum limits for ∆t. The Level 2 boundary for Flight Phase
Category C was reduced on the basis of the Concorde data in figure 51 which applies to the landing case
with pitch damper OFF. The C-5A Flight Phase Category C data in that figure, also for damper OFF, tend
to substantiate the proposed boundaries. In AFFDL-TR-75-3, however, the authors claim the C-5A should
be considered Level 1 with dampers OFF in the landing Flight Phase and, therefore, the Level 1 boundary
should be lowered. This recommendation was not accepted in preparation of the design criteria because
other data in AFFDL-TR-72-41 and AFFDL-TR-69-72 substantiate the higher Level 1 boundary. The
experiments reported in AFWAL-TR-81-3116 and AFFDL-TR-78-122 are the primary sources of data
suitable for establishing pitch limits for t1, the effective time delay parameter. The data in these reports is
given primary emphasis for Category C Flight Phases because the evaluation task was centered on
terminal Flight Phases including flare and touchdown. Data from the experiments reported in AFFDL-TR68-90 and AFFDL-TR-70-74 will also be used for Category A; these tasks were up-and-away or did not
include the critical flare and touchdown part of landing.
The correlation of pilot ratings for AFFDL-TR-78-122 is shown on figure 52, and the data from AFWALTR-81-3116 is shown on figure 52. In figure 52, the points at PR = 5, 6, 7, for t1 < 0.1 are not considered
in the data correlation because the pilot comments indicate that these configurations were downrated for
other reasons. The remaining data in figure 52 and the data in figure 53 indicate a rapid degradation in
pilot rating for the flare and touchdown task as t1 becomes greater than 0.1 sec. The data from the two
independent experiments are quite consistent and have been used as the basis for the pitch design
criteria. Although a band of values of t1 and PR is indicated by the data, nominal values of t1 have been
stated for the pitch design criteria.
As part of the experiment reported in AFFDL-TR-78-122, the evaluation pilots were requested to give
separate ratings for the approach and for the flare and touchdown. In addition, some approaches were
terminated by a waveoff and were rated for that task. It was found that the pilot ratings were significantly
less critical for the approach task than for the more critical flare and touchdown. The data on figure 54 are
pilot ratings for the approach segment or for the approach and waveoff. For this less critical task,
considerably larger values of t1 are tolerable.
229
MIL-STD-1797A
APPENDIX A
FIGURE 50. Comparison of YF-12 and XB-70 handing qualities evaluation with the
GPAS results (NASA-CR-159236).
230
MIL-STD-1797A
APPENDIX A
FIGURE 51. Short-period frequency (NASA-CR-159236).
231
MIL-STD-1797A
APPENDIX A
FIGURE 52. Pilot rating correlation with
effective time delay (AFFDL-TR-78-122
data).
FIGURE 53. Pilot rating with effective time
delay (AFWAL-TR-81-3116 data).
FIGURE 54. Pilot rating correlation with
effective time delay (AFFDL-TR-78-122).
FIGURE 55. Pilot rating correlation with
effective time delay (AFFDL-TR-68-90
data).
232
MIL-STD-1797A
APPENDIX A
The AFFDL-TR-68-90 experiment investigated the effects of higher-order control system dynamics on the
flying qualities for up-and-away fighter maneuvering and for landing approach to a low-altitude waveoff.
The data, plotted on figure 55, indicate that quite large values of t1 were on occasion considered flyable.
Review of the pilot comments for the cases with t1 > .35 indicate that the pilots were aware of a strong
pilot-induced oscillation tendency when they attempted precision closed-loop control or aggressive
maneuvering. Their ratings are all unacceptable for t1 > .35; the question being decided was whether or
not the airplane was controllable for rather undemanding tasks, i.e., they were not attempting air-air
combat maneuvering or gunnery tracking or actual landings. In addition, the cases not rated PR = 10
were generally ones for which the pilot had selected a low command gain which helped to reduce the
tendency for divergent PIO. The experiment included cases with a very slow feel system. Pilots
commented that they could recognize that the feel system was poor, i.e., they could individually sense
both the force applied and the stick motion; they tended to be more tolerant of these cases even though
large values of t1 characterize the response to a stick force step command. The major reason that
configurations with larger t1 (beyond the lines transferred from figure 52 and 53) were rated more
acceptable is thought to be that the evaluations were for a less critical task. This is certainly true for the
LA Group which was evaluated for the landing approach and waveoff. It may also be true for the A, B,
and C Groups even though they were supposedly evaluated for up-and-away flight in a fighter mission.
Review of the pilot comments indicates the evaluation task did not emphasize aggressive maneuvering
and tracking to the same extent as in AFFDL-TR-70-74.
Time history responses suitable to accurately measure t1 were not readily available for all of the
configurations evaluated in AFFDL-TR-70-74, but the data for Groups 1, 2 and 6 in that experiment are
plotted on figure 56. These data indicate a lower tolerance for effective time delay than was obtained from
the fighter up-and-away evaluations of AFFDL-TR-68-90. The data also indicate a reduced tolerance
relative to the data from AFWAL-TR-81-3116 and AFFDL-TR-78-122 which included landing flare and
touchdown. It is believed that the reduced tolerance to t1 indicated on figure 56 is a result of the emphasis
put on evaluation for air-air combat maneuvering and tracking capability in the AFFDL-TR-70-74
experiment.
Although the evaluation task has been introduced as a significant factor in the discussion of the data from
various experiments, the design criteria for effective time delay in pitch have not been stated in terms of
Flight Phases. The limits stated in the design criteria should be adequate to permit performance of flare
and touchdown during landing, which is probably the most critical maneuver in a transport mission, and
also for fighter maneuvering. To permit larger values for less critical flight phases is likely to invite too
casual an attitude toward this design problem (which we have on several occasions found to be poorly
appreciated by the design community) but is potentially the cause of very severe control problems such
as divergent PIO near the ground.
These requirements were written for a large supersonic-cruise, transport airplane (NASA-CR-159236,
which has been quoted/adapted extensively here), but the discussion relates the rise time to CAP and the
transient peak ratio to the damping ratio. Thus the requirements for other classes of aircraft can also be
converted from frequency to time domain. Since the presently recommended equivalent time delay limits
apply across the board to all Classes, the effective time delay measured from the step response would
also apply to all Classes. Note that this requirement retains the time-domain CAP [equal to (FS/n) MFS]
requirement as in the frequency-domain form, relating to attitude response.
Step control inputs do excite aircraft response at all frequencies. On the other hand they emphasize the
high frequencies and the steady state rather than the mid-frequency range of likely crossover, which is
critical for closed-loop pilot control. Such open-loop tests could not be expected to elicit all possible
closed-loop control difficulties. Time-domain measures may be more directly applicable to higher-order
and nonlinear responses, but step control inputs may be severely limited in size by the magnitude of
some responses, be
233
MIL-STD-1797A
APPENDIX A
they linear or nonlinear. Of course larger impulse or doublet inputs can be used, and the time-domain
requirements cast in such form.
For Level 3, see the discussion in section A and the supporting data in section B.
FIGURE 56. Pilot rating correlation with effective time delay (AFFDL-TR-70-74 data).
234
MIL-STD-1797A
APPENDIX A
D. Bandwidth, Time Delay
The bandwidth of the open-loop pitch attitude response to pilot control force for force controllers (pilot
controller force commands the control effectors) and to pilot controller deflection for deflection controllers
(pilot controller deflection commands the control effectors) shall be within the bounds shown on figure 57,
where ωBW is the highest frequency at which the responses of aircraft pitch attitude to pilot control-force
and control-deflection inputs have both 45 degrees or more of phase margin and 6 dB or more of gain
margin and
τp = -(φ2ω
180
+ 180°)/(57.3 x 2ω180)
where ω180 is the frequency corresponding to -180 deg phase and φ2ω
180
frequency.
is the phase angle at twice that
For Level 3, T2 , the time to double amplitude based on the value of the unstable root, shall be no less
than 6 seconds. In the presence of one or more other Level 3 flying qualities, ζsp shall be at least 0.05
unless flight safety is otherwise demonstrated to the satisfaction of the procuring activity. T2 applies to the
value of an unstable first-order root: T2 = -(ln 2)/λ where λ is the value of the unstable root.
FIGURE 57. Bandwidth requirements.
BACKGROUND
A measure of the handling qualities of an aircraft is its stability margin when operated in a closed-loop
compensatory tracking task. We refer to the maximum frequency at which such closed-loop tracking can
take place without threatening stability as bandwidth (ωBW). It follows that aircraft capable of operating at a
large enough value of bandwidth will have superior performance when regulating against disturbances. A
bandwidth criterion is especially useful for highly augmented aircraft in which the response characteristics
are non-classical in form (i.e., have large mismatch in equivalent system fits). Although not restricted to
such cases, this requirement should be utilized when the mismatch between the lower-order and higherorder systems exceed the values defined on figure 15. No assumption of pilot dynamics is necessary in
applying this requirement, since any such assumption would simply shift the boundaries. Also, for Level 1
minimal pilot compensation should be necessary.
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APPENDIX A
The concept of using bandwidth is not new for flying qualities. A 1970 utilization of bandwidth was in the
Neal-Smith criterion (see AFFDL-TR-70-74) consisting of empirical bounds on the closed-loop pitch
attitude resonance |θ/θc|max vs.- pilot equalization for a piloted closure designed to achieve a specified
bandwidth. Experience with this criterion has shown that the results can be sensitive to the selected value
of closed-loop bandwidth. The criterion developed herein bounds the value of bandwidth achievable
without threatening stability, thereby removing the necessity for selecting a value for ωBW a priori.
Another criterion utilizing bandwidth, suggested in AFFDL-TR-73-76, also selected a fixed value of
bandwidth (1 rad/sec for power approach). It utilized the phase margin, φM, and slope of the phase curve,
dφ/dω, at the selected bandwidth frequency as correlating parameters. Again, experience has shown that
the fixed value of bandwidth limits application of the criterion.
Most, if not all familiar handling quality metrics, are in fact related to bandwidth. However, these metrics
are generally tailored to classical aircraft which can be characterized by lower-order systems -- for
example, the q/FS and nz/FS transfer functions of section A, “CAP, …”.
It is easily shown for these (and similar) transfer functions that the quality of closed-loop error regulation
depends on the pilot's ability to increase the short-period root (ωsp) without driving it into the right half
(unstable) plane. As illustrated by the generic sketches in figure 58 for an idealized pilot supplying only
gain and pure delay, aircraft with low short-period damping ratio (ζsp), frequency (ωsp), or both, tend to
become unstable at low values of frequency (compare figures 58a and 58b). The aircraft of figure 58 is
represented as a simple short-period vehicle to simplify the example; for real highly augmented aircraft,
many more roots are involved.
Consider the bandwidth frequency as occurring at some (for now) arbitrary margin below the frequency of
instability (see boxes on root locus in figure 58). It can be seen from figure 58 that ωBW depends uniquely
on ωsp, ζsp, 1/Τθ2 , and τe (the delay, τe, draws the locus to the right as gain increases). Hence these
familiar flying quality metrics are, in fact, a measure of bandwidth. Again, we see that the flying qualities
application of bandwidth has roots in familiar metrics.
The present impetus for using ωBW as a criterion evolved from attempts to develop a flying quality
specification for aircraft utilizing unconventional response modes with direct force controls (wings-level
turns, pitch pointing, etc.), AFWAL-TR-91-3027. The infinite variety of responses that could occur due to
coupling within and between axes made it necessary to retreat to a more fundamental metric, which
turned out to be bandwidth. Strictly speaking, bandwidth in pitch involved θ/θC, a closed-loop describing
function of pilot/vehicle response. Here, however, a pilot model is merely a unity-gain feedback and so
bandwidth is specified in terms of the aircraft-alone gain and phase margins, only.
BANDWIDTH DEFINITION
The bandwidth as defined for handling quality criterion purposes is the highest frequency at which the
phase margin is at least 45 deg and the gain margin is at least 6 dB; both criteria must be met (figure 59).
Referring to figure 59, this describes the pilot's ability to double his gain or to add a time delay or phase
lag without causing an instability (Φ ≤ -180 deg at the ω for 0 dB gain indicates instability). In order to
apply this definition, one first determines the frequency for neutral stability, ω180, from the phase portion of
the Bode plot. The next step is to note the frequency at which the phase margin is 45 deg, ω135. This is
the bandwidth frequency as defined by phase, ωBWphase. Finally, note the amplitude corresponding to ω180
and add 6 dB. Find the frequency at which this value of response magnitude occurs; call it ωBWgain. The
bandwidth, ωBW, is the lesser of ωBWphase and ωBWgain.
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APPENDIX A
FIGURE 58. Simplified pilot-vehicle closure for pitch control.
If ωBW = ωBWgain, the system is said to be gain-margin limited; that is, the aircraft is driven to neutral
stability when the pilot increases his gain by 6 dB (a factor of 2). Gain-margin-limited aircraft may have a
great deal of phase margin, ΦM, but then increasing the gain slightly causes a large decrease in ΦM. Such
systems are characterized by frequency-response amplitude plots that are flat, combined with phase plots
that roll off rapidly, such as shown in figure 59.
Several sets of data were correlated with bandwidth using the above definition. A typical result is shown
in figure 60 utilizing the data from AFFDL-TR-70-74. While there is a definite pilot rating trend with ωBW,
the scatter for bandwidths between 2 and 6 rad/sec does not allow a quantitative definition of flying quality
levels. A detailed analysis of the pilot/vehicle closure characteristics was made for Configurations 1D and
2I, to determine why these two configurations with nearly equal ωBW would have such a large difference in
pilot ratings (4 and 8 respectively). The detailed pilot/vehicle closures are shown in figures 61a and 61b.
The value of bandwidth is seen to be about the same for both cases. However, if the pilot were to track
very aggressively by further increasing his gain he could increase the bandwidth of configuration 1D
greatly
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APPENDIX A
without adding much pilot compensation. Configuration 2I, though, offers a much more modest possibility)
of increasing bandwidth (compare the root loci in figures 61a and 61b). This behavior is predictable from
the phase curves: configuration 1D has a phase curve that rolls off very gradually at frequencies above
ωBW, whereas the phase for 2I drops off rapidly as the frequency is increased above ωBW. It is not
surprising that this case (2I) received a poor pilot rating (PR = 8) considering that attempts at aggressive
tracking result in a closed-loop divergence. A steeper phase dropoff means less ability to increase
bandwidth unless lead is added, or less increase in bandwidth for a given increase in pilot lead. Hence we
have evidence that the ability of the pilot to attain good closed-loop regulation without threatening stability
depends not only on
1) The value of bandwidth, ωBW,
but also on
2) The shape of the phase curve at frequencies above ωBW.
Rapid rolloffs in phase are well represented by a pure time delay, e-jωτ. Since that represents a phase
contribution of just -ωτ, both of the key factors noted above will be accounted for by plotting pilot rating
data on a grid Of ωBW vs. τ. This is done for the AFFDL-TR-70-74 data (which were plotted versus ωBW
alone in figure 60) as shown on figure 62. The scatter is seen to be considerably reduced and the data
are reasonably well separated into Level 1, 2, and 3 regions. The values of τ used in this plot were
obtained from lower-order equivalent system fits of the higher-order system transfer functions (MDC Rpt
A6792) The lower-order equivalent system form was:
θ
Fes
=
(
)
K θ s + 1 Tθ2 e − τes
[
s s 2 + 2ζ e ω e s + ω 2e
]
The zero 1/Tθ2 was fixed at the aircraft value (see earlier discussion of equivalent systems). But the
bandwidth criterion is intended to avoid the need for an equivalent system match. A workable and much
simpler approach is to note that to the extent that the rolloff in θ/FS phase beyond -180 deg can be
attributed to τe, we can estimate τe in the vicinity of some higher frequency ω1 (and associated phase φ1),
from:
τp =
− φ1 − 180 o
57.3 ω1
where ω1 is some frequency greater than the frequency for neutral stability3 and the symbol τp represents
the estimate of τe . Correlations between τe and τp for the combined AFFDL-TR-70-74 and AFFDL-TR-78122 data resulted in a correlation coefficient of 0.96. Thus, there is very good evidence that τp can be
used in place of τe in figure 62, as will be shown in Supporting data.
3
ω1 was taken as twice the neutral stability frequency, i.e. ω1 = 2ω180. Hence τ = -(φ2ω
180
+ 180°)/(57.3 x
2ω180). At the frequency 2ω180, structural or other modes may complicate analysis, thus making
application of this requirement difficult.
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APPENDIX A
FIGURE 59. Definition of bandwidth frequency ωBW from open loop frequency response.
FIGURE 60. Comparison of Neal-Smith data (AFFDL-TR-70-74) with bandwidth (mean ratings).
239
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APPENDIX A
240
MIL-STD-1797A
APPENDIX A
FIGURE 61a. Level 1/2 system of Neal-Smith (lD): ωBW
= 2.7 rad/sec, mean PR = 4.1.
FIGURE 61b. Level 3 system of Neal-Smith (2I) ωBW =
2.5 mean PR = 8.0.
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APPENDIX A
FIGURE 62. Correlation of pilot ratings with ωBW and τe (AFFDL-TR-70-74 data).
Responses that are gain-margin-limited tend to have shelf-like amplitude plots as shown on figure 63.
With such systems a small increase in pilot gain results in a large change in crossover frequency and a
corresponding rapid decrease in phase margin. The decrease in phase margin becomes critical for
attitude control when τp is moderately large (of order 0.1 to 0.2). The two configurations shown on figure
63 are taken from the AFFDL-TR-78-122 experiment. Applying the previously discussed definition of
bandwidth, we find that both Configurations 5-6 and 5-7 are pin-margin-limited. Both configurations suffer
from the same deficiency, i.e., moderate values of τe combined with a shelf-like amplitude curve that
results in a very rapid decrease in phase margin with small changes in pilot gain. However, the 6 dB limit
selected to define ωBWgain does not catch Configuration 5-6. While this configuration is correctly predicted
to be Level 2 (PR = 6) on the basis of τp, the value of ωBW is in the Level 1 region. Had a slightly higher
value of gain margin been picked to define ωBW, the bandwidths for Configurations 5-6 and 5-7 would be
approximately equal. However, because of the nature of shelf-like frequency responses, there will always
be a case which can fool the criterion. An experienced handling qualities engineer would immediately
recognize the shelf-like shape and moderate τp as a significant deficiency. However, the purpose of a
criterion is to eliminate such judgement calls. Nonetheless, it is not expected that this idiosyncrasy will
result in problems with correlating or predicting pilot rating data inasmuch as moderate (Level 2) values of
τp are required to get misleading values of ωBW (i.e., rapid phase rolloff in a frequency region where the
amplitude curve is flat must occur to get the effect shown on figure 63).
Supporting data
The data from Neal-Smith (AFFDL-TR-70-74) are compared with the bandwidth Category A requirements
on figure 64. Some points with discrepancies between the rating and the LOES criteria [table
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MIL-STD-1797A
APPENDIX A
XIII] are filled in on figure 64. The bandwidth criteria correctly evaluate three or four of these five points.
These results are reasonably encouraging, though there are a number of Level 2 ratings at high values of
bandwidth. The abbreviated pilot comments (taken from AFFDL-TR-70-74 and AFFDL-TR-74-9) indicate
that abruptness and oversensitivity become a problem when ωBW is large. This was especially true of the
AFFDL-TR-74-9 pilot ratings (given in parentheses on figure 64). A possible boundary on ωBW is shown
on figure 64 to account for this problem. This boundary is considered tentative because the issue of
overresponsiveness is not completely understood at this time. A broader data base is felt to be necessary
to verify the results concerning an upper limit on ωBW, so this is indicated by a broken line on figure 57.
FIGURE 63. Large difference in bandwidth due to shelf in amplitude plot with
moderate values of τp (configurations of AFFDL-TR-78-122).
243
MIL-STD-1797A
APPENDIX A
FIGURE 64. Correlation of pilot ratings with ωBW and τp for Neal-Smith data (Category A) (data
from AFFDL-TR-70-74, ratings in parentheses from AFFDL-TR-74-9).
The evaluation maneuvers performed in the Neal-Smith study included a pitch-bar tracking test but did
not have an actual air-to-air tracking task. When tracking a target aircraft there is some suggestion of
acceptance of abruptness . For example, Configuration 13 in the AFFDL-TR-70-74 experiments was
rated 7 and 5.5 due to excessive sensitivity. However, in a follow-on experiment (AFFDL-TR-74-9) with a
target aircraft, Configuration 13 was rated a 2 on two separate evaluations. At first glance this would
seem to be an idiosyncrasy of different pilots in a different experiment; but during a repeat experiment,
.
the target aircraft was removed and the rating went from 2 back up to 7 (see ◊ on figure 64).
The data correlations on figure 64 represent up-and-away flight, and so are appropriate for generating
boundaries for Category A. Data (AFFDL-TR-78-122) for Category C (approach and landing) are
correlated with ωBW and τp on figure 65. The upper boundary on (ωBW for Level 1 is considered tentative
for the reasons discussed above.
The bandwidth criterion was developed for highly augmented aircraft, and the data shown in its support
have been for high-order systems. Figure 66 and 67 compare bandwidths of classical (unaugmented)
airplanes with pilot ratings obtained in flight simulations. For AFFDL-TR-66-63, AFFDL-TR-68-91,
Princeton University Rpt 777, and AFFDL-TR-69-3, the test vehicle was the USAF/CALSPAN T-33, for
which τp = 0.07 sec (due to actuation and feel systems); τp for the NASA-TN-D-3971 data, a Boeing 36780, is not known but is assumed to be about the same.
The classical-airplane data agree rather well with the Level 2 and 3 boundaries, but for both Categories A
and C the Level 1 boundary of figure 57 appears too stringent. (For example, in Category A flight, figure
57a does not allow τp greater than about 0.06 sec for Level 1, therefore all the figure 65 data (for τp =
0.07) should be rated Level 2 or worse. The data, however, tend to support a Level 1 boundary at ωBW = 4
rad/sec, as shown by the dashed line. No rating worse than 4-1/2 was given for ωBW > 4. The reasons for
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MIL-STD-1797A
APPENDIX A
this disagreement have not been resolved, though the task used for evaluation, as discussed earlier, may
not have been tight enough to provoke pilot objections to response abruptness or to excessive time
delays.
No supporting data are available at this time to establish Category B boundaries.
Somewhat unstable configurations, with no bandwidth at all, can be flown quite safely (see discussion of
Level 3 requirements in Supporting data). Therefore, for statically unstable aircraft the Level 3
requirement stated earlier should be applied.
Reaction to these bandwidth criteria for pitch response has been mixed. While some have had moderate
or even better success with it, others comment that a) the criteria do not predict flying qualities Levels
correctly; b) the bandwidth criteria exclude some of the Level 1 CAP area, and vice versa; c) there are
cases in which τp is not close to τe, and d) path control is not addressed. Nevertheless, because of the
success cited we recommend using bandwidth along with other criteria.
For Level 3, see the discussion in section A and the supporting data in section B.
FIGURE 65. Correlation of pilot ratings with ωBW and τp for approach and landing (AFFDL-TR-122).
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MIL-STD-1797A
APPENDIX A
247
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APPENDIX A
FIGURE 66. Comparison of pilot ratings for Category A short-period configurations with bandwidth (classical airplanes).
248
MIL-STD-1797A
APPENDIX A
FIGURE 67. Comparison of pilot ratings for Category C short-period configurations
with bandwidth (classical airplanes).
249
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APPENDIX A
E. Closed-Loop Criterion
For pilot control of pitch attitude in the manner sketched,
where Yc is the transfer function of the aircraft and flight control system, θc is an external pitch command,
θe is pitch attitude error, F is pilot force on the pitch controller, δ is its deflection and Yp is the analytical
pilot model (either form is permitted):
Yp = K p e −.25s
(Tp1s + 1)
(Tp2 s + 1)
or
K p e −.25s
(5s + 1)
s
(Tp1s + 1)
(Tp2 s + 1)
a bandwidth, defined by a closed-loop phase of -90 degrees, of
FLIGHT PHASE
BANDWIDTH
Category A
3.5 rad/sec
Category B
1.5 rad/sec
Landing
2.5 rad/sec
Other Category C
1.5 rad/sec
shall be attainable with closed-loop droop no more than -3 dB for Levels 1 and 2 and closed-loop
resonance no greater than 3 dB for Level 1, 9 dB for Level 2 over the frequency range from 0 to 10
rad/sec. The pilot model is constrained to the given forms but there are no limits on Kp, Tp1 or Tp2. The
requirements apply for both force and deflection pilot control inputs. Figure 68, in the form of a Nichols
chart, illustrates these limits. The pilot output is force for force controllers (pilot controller force commands
the control effectors) and deflection for deflection controllers (pilot controller deflection commands the
control effectors).
For Level 3, T2, the time to double amplitude based on the value of the unstable root, shall be no less
than 6 seconds. In the presence of any other Level 3 flying qualities, ζsp shall be at least 0.05 unless flight
safety is otherwise demonstrated to the satisfaction of the procuring activity. T2 applies to the value of the
unstable root: T2 = - (ln 2)/λ where λ is the value of the unstable root.
The criteria contained in this paragraph are intended to ensure good dynamic performance capability of
the pilot-aircraft dynamic system. The form in which the criteria are stated was selected to permit
accommodation of highly augmented aircraft and systems with transport time delay or cascaded dynamic
elements. Through application of describing function techniques, the criteria may also permit investigation
of the effects of certain nonlinearities typically encountered in flight control systems.
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MIL-STD-1797A
APPENDIX A
FIGURE 68. Design criteria for pitch dynamics with the pilot in the loop.
251
MIL-STD-1797A
APPENDIX A
The criteria as modified by Chalk (NASA-CR-159236) are derived from the work by Neal and Smith
reported in AFFDL-TR-70-74. The basic approach is to model the pilot-aircraft pitch attitude control loop
as a unity-feedback system with a pilot model of an assumed form in the forward loop. The form of the
assumed pilot model permits accounting for the following characteristics exhibited by pilots when
controlling dynamic systems:
• Adjustable gain.
• Time Delay
• Ability to develop lead, or to operate on derivative or rate information.
• Ability to develop lag, or to smooth inputs
• Ability to provide low-frequency integration
Neal and Smith's original version also placed limits on phase compensation in the pilot model.
The two forms of the pilot model account for the observed capabilities and limitations of the pilot with
sufficient accuracy to permit approximate analysis of the dynamics of the closed-loop pilot-aircraft system
in pitch. It should be emphasized that the pilot model need not be an exact analog of the human pilot in
order to be useful in the context of design criteria. The criteria are based on the hypothesis that if good
closed-loop dynamic performance can be achieved with an autopilot with the characteristics described by
the assumed pilot model, then the human pilot will also be able to achieve good closed-loop dynamic
performance with acceptable workload.
These two forms differ by the (5s + 1)/s term, the low-frequency integration capability. It is intended that
the form of the model without this term will be used when constant speed or two-degree-of-freedom
equations are used to represent the aircraft. In this case the aircraft transfer function should have a free s
in the denominator and low-frequency integration by the pilot will not be necessary. When three-degreeof-freedom equations are used or when the flight control system uses attitude stabilization ("Aircraft
Dynamics and Automatic Controls"), it may be necessary for the pilot model to perform low-frequency
integration to avoid droop at frequencies less than ωBW.
The e-.25s term in the pilot model accounts for time delay in the pilot's neuromuscular system. The value of
0.25 sec. is based on delays observed in records for the discrete tracking task performed in AFFDL-TR70-74 and AFFDL-TR-78-122. These records exhibit delays ranging from 0.20 to 0.40 seconds. The value
of 0.25, selected on the basis of cut-and-try data correlation, is interrelated with the bandwidth frequency
that is specified for a given flight phase or task. The values of time delay (0.25 sec) and bandwidth (ωBW =
2.50 rad/sec) for the landing Flight Phase have been determined from empirical correlation of data in
AFFDL-TR-72-143 and AFFDL-TR-78-122.
Because the closed-loop pilot-aircraft dynamic system has been modeled as a negative feedback system
with unity gain in the feedback path, it is possible to relate the dynamic characteristics of the elements in
the forward loop, θ/θe = YpYc, to the dynamic characteristics of the closed-loop system, θ/θc = YpYc/(1 +
YpYc), through use of a Nichols chart. This diagram consists of the superposition of two grid systems. The
rectangular grid is the magnitude and phase of the forward-loop dynamic system, YpYc, while the curved
grid system represents the magnitude and phase of the closed-loop system, θ/θc = YpYc/(l + YpYc).
Therefore one can determine the closed-loop dynamic characteristics by plotting the magnitude and
phase of YpYc over a range of frequency on the rectangular grid.
It is hypothesized that a given Flight Phase or task performed in a typical environment will require certain
minimum dynamic characteristics of the closed-loop pilot-aircraft system. The parameters used to define
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APPENDIX A
the closed loop dynamic performance are bandwidth, droop at frequencies below the bandwidth and
resonance magnitude. These closed-loop system parameters are defined by the curved lines on figure
68. The maximum droop permitted for ω < ωBW is -3.0 dB. This value has been defined somewhat
arbitrarily but can be justified from examination of discrete tracking task records in AFFDL-TR-70-74 and
AFFDL- TR-78-122 and by interpretation of pilot comments in these documents.
Hanke (AGARD-CP-333) relates the Neal-Smith criteria to ω 2sp /(n/α).
SUPPORTING DATA
The bandwidth frequency and the closed-loop system resonance limits for Level 1 and Level 2 have been
determined from empirical data correlation, AFFDL-TR-70-74 and Radford and Smith in AFWAL-TR-803067.
It is not feasible to present all the data available in the literature as substantiation of these closed-loop
criteria. Four configurations have been selected for presentation to illustrate characteristics of interest to
designers.
NASA-CR-159236 selected configurations 12 and 13 from AFFDL-TR-72-143 to illustrate dynamics
typical of these aircraft that might result from center of gravity variations. Configuration 13 is
representative of a forward c.g. while Configuration 12 represents an extreme aft c.g. which results in an
unstable real root with a time to double amplitude T2 = 2.1 sec.
The YpYc data for these two cases are plotted on figures 69 and 70. From the plot on figure 69 it is seen
that it was possible to find Kp, Tp1 values that would satisfy the design criteria for Level 1 for Configuration
13. This configuration was rated PR = 1-1/2, 2, 3 on three separate evaluations. The plot on figure 69
indicates, that it was not possible to find pilot compensation that would satisfy the Level 1 design criteria
for Configuration 12. It was possible, however, to satisfy the Level 2 design criteria. The pilot ratings for
Configuration 12, however, were PR = 10. To understand this rating, it is necessary to realize that to
obtain the plot illustrated, the pilot would have to develop large values of Tp1 and a very specific value of
Kp. This means he must operate on pitch rate information and he must constantly close the control loop
with very little variation in the pilot model parameters. Thus in reality there are more dimensions to the
criteria than are explicitly indicated by this requirement. [A good aircraft is one for which the closed-loop
performance is not critically dependent on the values of the pilot model parameters.) The open-loop
divergence limit of 4.4.1 on time to double amplitude would identify Configuration 12 as an unacceptable
design, i.e. the behavior of this aircraft with no pilot control is unacceptable.
Configurations 6-1 and 6-2 from AFFDL-TR-78-122 were selected in NASA-CR-159236 to illustrate the
detrimental effects that can result from cascading dynamic elements in the flight control system.
Configuration 6-1 had a second-order prefilter with ωn = 4 rad/sec and ζ = .7. Attempts to land this
configuration resulted in uncontrollable pilot-induced oscillations; pilot rating was 10. Removal of this
second-order prefilter and substitution of a lead-lag prefilter, (.06s + 1)/(.10s + 1), reduced the phase shift
and resulted in greatly improved flying qualities which were given a pilot rating of PR = 2. Figure 71
illustrates that Configuration 6-1 would exhibit large resonance when compensated by the pilot to achieve
ωBW = 2.5 rad/sec without violating the -3 dB droop constraint. Figure 72, the plot for Configuration 6-2,
illustrates that the Level 1 design criteria can easily be met by adjustment of the pilot model parameters,
for a range of values of Kp and Tp1; thus the closed-loop performance of Configuration 6-2 is not critically
dependent on specific values of the pilot model parameters. Pilot rating was 2.
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APPENDIX A
FIGURE 69. Amplitude-phase plot for
configuration 13 (fwd; c.g.).
FIGURE 70. Amplitude-phase plot for
configuration 12 (aft; c.g. - unstable).
255
MIL-STD-1797A
APPENDIX A
FIGURE 71. Amplitude-phase plot for
configuration 6-1.
FIGURE 72. Amplitude-phase plot for
configuration 6-2.
256
MIL-STD-1797A
APPENDIX A
Configurations 6-1 and 6-2 illustrate the degradation in flying qualities that ran result from excessive
phase shift in the pilot's command channel.
Although the Neal-Smith criteria account explicitly for pilot compensation, they may not always
adequately account for sensitivity of closed-loop performance to the pilot-chosen bandwidth (see Radford
and Smith in AFWAL-TR-80-3067).
Also note that the Neal-Smith criteria do not address path control at all; they examine only the pitch
attitude loop. Additional criteria would seem necessary to assure adequate path control: perhaps
ω 2sp /(n/α) or ωsp Tθ2. Sarrafian (NASA-TM-86728) has used another variant of the Neal-Smith technique
to correlate approach and landing data from two TIFS variable-stability airplane evaluation programs. He
closed an inner pitch attitude loop, for all configurations, with the same amount of pilot phase
compensation. Then closing an outer flight-path loop, he found that the achievable bandwidth correlated
the pilot ratings.
For Level 3, see the discussion in section A and the supporting data in section B.
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APPENDIX A
F. Time- and frequency-response criteria by Gibson
The criteria based on the step input time response features shown on figure 73 are primarily related to
attitude and flight path. They are intended for fly-by-wire control law design optimization and overall
handling Levels 1, 2, and 3 have not been established.
a. For Category A and Category C Flight Phases, attitude dropback as defined on figure 73
should not normally be negative. Satisfactory values depend on the task and on the pitch rate transients.
b. Normal acceleration responses can be related to Level 1 frequency and damping
requirements by the boundaries shown on figure 74. Any oscillations following the first peak should
subside such that the ratio of successive half-cycles is less than 0.3.
c. Boundaries of satisfactory frequency responses for Category A precision attitude tracking are
shown on figure 75. Responses outside the low-frequency limits with the handling tendencies indicated
may be satisfactory for other tasks, as determined by time response limits.
d. An envelope of satisfactory Category C landing approach response for frequencies below the
required bandwidth of 0.25 to .5 Hz at 120 degrees phase lag is shown on figure 75.
e. All frequency responses must satisfy the figure 75 requirements for response attenuation and
phase lag rate of increase at the 180 degree phase lag crossover frequency.
The following discussion by Gibson is extracted from AGARD-CP-333 and ICAS-86-5.3.4. Figures 76 to
82 further illustrate these concepts and define his parameters.
USAF-Calspan in-flight simulations of pitch handling with high-order flight control systems, using their T33 aircraft, have furnished a source of information about acceptable attitude and flight path
characteristics. In AFFDL-TR-74-9 the tasks of air combat maneuvering, air-to-air tracking and flight
refueling, and in AFFDL-TR-78-122 the approach and landing tasks were assessed for wide ranges of
basic frequency and damping modified by stick prefilters to simulate high-order effects
Step responses were calculated for all these configurations and their features compared with pilot rating
and comment data. Some quite clear results were obtained which can be summarized as follows.
•
Negative attitude drop back (i.e. overshoot) was usually associated with sluggish, unpredictable
response both in flight path control and in tracking, leading sometimes to overdriving PIO
•
Attitude drop back from 0 to about 0.25 seconds was excellent for fine tracking and was
associated with comments typified by the nose follows the stick.
•
Increasing attitude drop back with large pitch rate overshoot led to abrupt response and bobbling,
from slight tendency to continuous oscillations, in tracking tasks. Sometimes this was called PIO,
but it did not cause concern for safety.
•
Attitude drop back had little effect within the range tested upon gross maneuvering without target,
landing approach or flight refueling, provided it was not negative.
•
CAP up to 3.6 rad/sec2/g was satisfactory for gross maneuvering without a target, but was
unsatisfactory above 2 rad/sec2/g for the landing approach, above 1 rad/sec2/g for fine tracking,
and below 0.28 rad/sec2/g for any task.
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FIGURE 73. Pitch short period time responses.
259
MIL-STD-1797A
APPENDIX A
FIGURE 74. Equivalent ωn boundaries.
FIGURE 75. Design aim criteria for pitch attitude frequency response.
260
MIL-STD-1797A
APPENDIX A
FIGURE 76. High and low order frequency response.
FIGURE 77. Trends of high order phase rate.
261
MIL-STD-1797A
APPENDIX A
FIGURE 78. Flight path - attitude relationships.
FIGURE 79. nz hang-on effects.
262
MIL-STD-1797A
APPENDIX A
FIGURE 80. Precision tracking: q ∼ θ trends.
FIGURE 81. New short-period thumbprint (from ICAS-86-5.3.4).
263
MIL-STD-1797A
APPENDIX A
•
FIGURE 82. Low-order pilot - aircraft attitude frequency response.
The pitch rate overshoot ratio seems to qualify the drop back behavior, with a value greater than
3.0 resulting in unacceptable drop back as small as 0.25 seconds. The trends are indicated on
figure 80.
•
Small values of flight path time delay were associated with excellent flight refueling control, but
were not essential for good gross maneuvering and did not on their own ensure predictable
behavior.
•
Overshoots in normal acceleration did not cause unpredictable behavior unless associated with
low frequency, breaching the upper right-hand time response boundary of figure 74.
These factors point to a number of design criteria and trend indicators which have been put to use in
control law developments with excellent results. No attempt has been made to define the equivalent of
Level 1, 2 or 3 limits and the most appropriate location seems to be here in the handbook, as guidance.
One result stands out clearly as a candidate for a new requirement. This is to specify that attitude drop
back be zero or positive, as negative values were always rated sluggish and unsatisfactory in these
experiments. This leads also to the result that some pitch rate overshoot is always necessary for optimum
handling, with the function of minimizing drop back for tracking inputs or of rapid generation of the angle
of attack increment required for crisp flight path response in gross maneuvers, landing flare, etc. It has no
significance otherwise for the pilot unless it becomes too large. However, it has been shown in other
results that a small attitude overshoot can be satisfactory for basic attitude control, e.g. in aerobatics,
cruise, etc., and for attitude tracking with rigid or small travel sticks.
The principal time response features can be identified in the ωsp Tθ2 - ζsp format discussed earlier, as 4D
shown on figure 81. Some are constant for all Tθ2 and the others vary in proportion to Tθ2.
264
MIL-STD-1797A
APPENDIX A
Constant features:
•
Zero attitude dropback line, along which the flight path time delay tγ equals Tθ2.
•
Pitch rate overshoot ratios.
•
PIO line where 2ζsp ωsp = 1/Tθ2, an artifact of variable stability experiments not achievable by
aerodynamic means. Serious PIO and extreme pilot nonlinearity occur below this boundary.
Variable features:
•
Times to the first pitch rate peak, tq.
•
Non-zero attitude dropback, along which tγ = (Tθ2 - dropback). Note tγ = 2ζsp/ωsp also, figure 78.
Analysis of low-order aircraft indicate that the satisfactory range of Category A short period frequency is
limited by values of tq between about 0.3 and 0.9 seconds. The area of satisfactory handling therefore lies
in the upper pan of the figure for large Tθ2 and moves down to the lower part with reducing Tθ2. Larger
values of tq are satisfactory for Category C, e.g., from 0.5 to 1.1 seconds for Class 4 aircraft up to possibly
2.0 seconds for large aircraft. With typical Tθ2 of 1.5 to 2.0 at low speed this is in a region of large
dropback and pitch rate overshoot, but these are acceptable up to at least 1.5 and 3.0 respectively.
The PIO line forms a lower limit which should not normally be approached if conventional damping
augmentation practices are followed, a practical limit slightly above it being expected. The zero dropback
line should normally be respected but in some cases the limit may be above it, e.g. in the landing
approach with Tθ2 = 2.0 the satisfactory Class 4 tγ limit of 1.5 seconds sets a minimum of 0.5 dropback.
With smaller Tθ2 a lower frequency can make a nominal zero dropback look like a large overshoot in the
first few seconds of the landing flare, but this is avoided by observing the tq limit.
These time response carpets cannot be used for plotting high-order aircraft results, as they will contain a
mixture of aerodynamic and FCS modes. Although the dropback and path delay remain connected by Tθ2,
they and the pitch rate overshoot and time to peak no longer have a unique relationship to ωspTθ2 and ζsp.
For the analysis or design of such systems the time-response features are considered individually.
Additional high-order effects will be evident most importantly in tq., the delay in reaching the pitch
acceleration peak which is a strong indicator of PIO and handling problems when greater than 0.3
seconds. The elimination of this defect is achieved by the frequency response shaping discussed later.
A conclusion to be drawn is that for low-order aircraft with elementary pitch damper augmentation, a low
maneuver margin should be aimed for, with its inherently high natural damping, if precision pitch handling
is required. This is completely consistent with the excellent Lightning low-altitude, high-speed (LAHS)
pitch handling characteristics where in fact a frequency lower than the MIL-F-8785C minimum is
satisfactory, together with only 2.0 lb/g stick forces. The much larger TSR2 prototype also had a low
maneuver margin, with good damping in the LAHS region and was taken on only its 20th flight to 550
knots at 250 feet over hilly terrain without any stability augmentation. It was rated as having control and
response well matched to this task.
From the combination of the facts that the MIL-F-8785C frequency is proportional to speed and the flight
is path time delay is inversely proportional to frequency (given constant maneuver margin and damping),
and hence that this time delay is inversely proportional to speed, it will be observed that the path distance
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MIL-STD-1797A
APPENDIX A
represented by the delay is constant. In effect this reveals that the flight path response bandwidth could
be considered to be constant and independent of speed, which may be of relevance to close-in air-to-air
combat. If this is the case then this result is compatible with the concept of a fixed attitude frequencyresponse bandwidth independent of speed also, a subject discussed here and also in AFWAL-TR-823081.
We note that pitch-attitude overshoot (negative drop back, figure 73) and normal-acceleration hang-off
tendencies are sometimes found in rate command/attitude hold systems. One solution then is to quicken
the sluggish response by addition of a feed forward directly to the control surface.
Much work was done in attempting to predict pilot opinion from closed-loop analyses, though this does
not seem to have been followed in more recent years with the exception of the Neal and Smith criterion.
With the advent of the computing power potential of digital fly-by-wire it is now much more useful to the
flight control system designer to attempt to shape the aircraft frequency response into a form known to be
attractive to the pilot, with which he can perform both well and easily and hence will result in a good
opinion rating. The pilot model which achieves this aim is well known to be the simple gain and time
delay, the latter always being present in random error tracking. It is possible to define an envelope of
aircraft attitude response which is very robust, in the sense that the pilot can achieve good closed-loop
control with a wide range of gain and delay only. In this approach it is unnecessary to define a pilotvehicle bandwidth since he has a wide choice according to the needs of the task.
The classical aircraft dynamics which have always been shown to achieve the best ratings in simple
tracking experiments are a pure-gain pitch rate response and the resulting attitude response of K/s. Real
aircraft have inertia, control power limits, and pilots who dislike excessive pitch acceleration, and can only
be represented by this model at low frequencies. These attitude responses are indicated on figure 82
using the Nichols' chart form on which open-loop and closed-loop responses are related. Because of this
facility these charts are often more useful to the FCS control law designer than the more usual Bode
plots, even where no pilot model is being added to the aircraft response. On figure 82 a pilot gain and
delay model is added to a pure low-order attitude response to show good closed-loop performance with
negligible droop or resonance. The gain is chosen to give the pilot-vehicle open-loop crossover frequency
of 0.3 Hz, and a small delay typical of simulation results is selected. A K/s response is included for
comparison with the aircraft response with the crossover frequency.
This basic pilot model assumption underlies the aircraft response boundary limits used as a design
criterion. The crossover frequency typifies the upper end of the 1 to 2 rad/sec. range and the 0.2 second
time delay typifies the pilot delay noted generally in the literature in simulation experiments. WADC-TR57-520 measured the difference between flight and simulation to show that, while the lead or lag
equalization did not change, the pilot gain was lower and the time delay was larger in flight.
Choice of these values therefore represents an upper limit on pilot performance in the definition of aircraft
response boundaries. The choice of frequency in Hz rather than radian/second is deliberate since the
pilot sees frequency behavior in terms of its period or cycles per second, and this serves to present a
more obvious view of the effect of such boundaries.
Figure 75 shows optimum aircraft pitch attitude response boundaries for precision tracking tasks, in which
the crossover frequency of 0.3 Hz is inherently achieved with a low pilot gain appropriate to an aircraft low
stick force per g. If this criterion is satisfied, the allowable pilot phase lag for optimum tracking can be
attributed to his time delay and no further equalization is required from him. If the pilot chooses a lower
crossover frequency, the allowable lag increases and he can adopt a larger time delay without departing
from a good closed-loop performance.
These boundaries do not represent overall Level 1 limits. Depending on the task, responses outside them
can be very satisfactory. General characteristics associated with areas outside the boundaries are
indicated,
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and were derived from correlation with comment data from AFFDL-TR-74-9 and up-and-away flight
configurations. While the boundaries represent the small number of cases which were optimum for all
tasks, some cases attained Level 1 ratings for flight refueling despite bobble severe enough to degrade
pitch tracking to Level 2, and the best flight refueling case (really excellent) was close to the boundary
confines but was a poor Level 2 for gross maneuver because of unpredictable g response. Despite this it
had excellent flight path control with small delay, provided that aggressive control was not attempted.
More generally, attenuation greater than 5 or 6 dB/octave near the crossover frequencies is associated
with sluggish, unpredictable flight path control.
PHASE LAG
For an ideal low-order aircraft the maximum possible phase lag is zero for pitch acceleration, 90 degrees
for pitch rate, 180 degrees for pitch attitude, 270 degrees for flight-path angle, etc. Long experience has
shown that the addition of moderate actuation phase lags need not alter the essential low-order
characteristics so far as the pilot can observe them. It is also well established that good handling qualities
are confined to regions within this broad definition of low-order systems.
Hence an overriding consideration for high-order flight control system design should be an attempt to
contain phase lags to values no greater than the above plus say an extra 30 degrees for all frequencies
below 1.5 Hz or preferably even 2.0 Hz.
OTHER CONSIDERATIONS
Gibson's views on pilot-induced oscillations are discussed under 4.2.2 Guidance.
For Level 3, see the discussion in section A and the supporting data in section B of 4.2.1.2 Guidance
REQUIREMENT LESSONS LEARNED
Three major lessons have emerged from recent work on equivalent systems, and from flight experience
with several prototype airplanes.
a. There are sufficient parameters in the equivalent system models to allow correlation with flying
qualities problems of the very high-order systems which have so far been designed for operational aircraft
b. Of these equivalent parameters, large equivalent delays are highly correlated with pilot-induced
oscillation tendencies.
c. Succumbing to the temptation to add complexity to the flight control system can easily degrade, rather
than improve, the handling qualities.
The second lesson, though evident in the in-flight simulation data of DiFranco, Neal and Smith, and
LAHOS (AFFDL-TR-68-90, AFFDL-TR-70-74, and AFFDL-TR-78-122 respectively) has also been learned
the hard way. The Tornado experience described by Gibson in AGARD-CP-333 was discussed in
AGARD-AR-134 as follows:
[The Tornado description] is a rare example of a type of paper that should be
encouraged. In this paper the airplane designer admits that his airplane, equipped though
it is with a full authority fly-by-wire flight control system, turned out to have serious flying
qualities problems that required solutions. The example is rare not because problems
occurred, but because the designer was willing to report on the experience. In fact,
similar problems (pitch PIO in landing caused by control system phase shift and roll PIO
caused by high roll control gain) have been experienced in highly augmented aircraft
designed in the USA such as the YF-17, YF-16, F-18, and Space Shuttle.
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MIL-STD-1797A
APPENDIX A
In the Tornado example, the problem was related to excessive pitch command pins and the phase lag
from low-frequency filters (i.e., large -τe). Richards and Pilcher (“SETP Cockpit”) give a frank discussion of
PIOs (lateral in this case) encountered when the demanding task of shipboard landing was first evaluated
with an early F-18 version containing excessive equivalent delay.
An important lesson learned from both the Tornado and F-18 experience is that the pilot-induced
oscillations due to equivalent delay, or phase shift, though pronounced, can be very isolated. Lateral PIOs
occurred in two of the 49 carrier landings performed with the F-18. Considerable flight experience had
been accumulated on the Tornado before the hard landing reported by Gibson.
Differences between ground-based simulation and in-flight characteristics appear inherent in experience
with the above aircraft. Presumably all these aircraft were simulated on ground equipment during the
design, and their problems only appeared later in flight. The differences, seen in the early results of
DiFranco (AFFDL-TR-68-90) and Parrag (AFFDL-TR-67-19) have also been the subject of some recent
study. Figure 83 illustrates some differences between pilot ratings for various equivalent delays in various
simulations. The figure is from AlAA Paper 80-1611-CP, which summarizes the lessons learned:
FIGURE 83. Comparison of equivalent delay effects in pitch or roll response to stick
force for different simulations (from AIAA Paper 80-1611-CP).
Pilot rating degradation due to equivalent delays is often far more serious in flight than on
a ground-based simulator.... Most of the data show a threshold in pilot rating degradation
due to delay followed by a fairly linear increase in the rating.
The Navion in-flight results [“Pilot Opinions for Sampling Effects in Lateral-Directional
Control" and AIAA Paper 79-1962] form both extremes of the data, i.e., producing the
most immediate degradation due to delay (for lateral dynamics) and also the least
ultimate degradation (for longitudinal dynamics).
The MCAIR ground-based data are similar to the F-8 low stress landings of Berry, et al
[AIAA Paper 80-1626-CP]. The F-8 high stress landing data closely approach the NT-33
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MIL-STD-1797A
APPENDIX A
longitudinal landing data [AFFDL-TR-78-122 and AFWAL-TR-81-3116] and the NT-33
lateral landing data [AIAA Paper 79-1783 and AFWAL-TR-81-3116]. A general trend of
rating versus delay can be inferred.... However, there is much to be learned about lags
and equivalent delay effects.
In the Supporting Data section an example of the value of equivalent systems was shown for an
augmented aircraft. A more recent application of LOES provides even stronger support. An emergency
backup control system, for the USAF AFTI/F-16 power approach and landing, was designed with a pitch
rate feedback to the horizontal tail. Figure 84 illustrates the control system, and shows the θ/δe transfer
function for the unaugmented AFTI/F-16 in the power approach (129 kt, 13.2 deg α). The feedback is
intended to stabilize the short period, which consists of two fist-order modes for the basic airplane. As
figure 85 shows, the augmented airplane has two well-damped second-order modes, and the short-period
2
) is well within Level 1 limits for damping ratio
mode (resulting from the coupled pitch rate lag and 1/ Tsp
and frequency.
FIGURE 84. AFTI/F-16 independent back-up pitch rate feedback block diagram.
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MIL-STD-1797A
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FIGURE 85. AFTI/F-16 q → δe feedback (IBU).
However, when this system was simulated and flown on the USAF/Calspan NT-33, it received CooperHarper pilot ratings of 8 and 9, and was considered extremely sluggish with very heavy control forces. An
equivalent system match of the θ/Fs transfer function (see figure 86) clearly shows why the airplane was
Level 3: equivalent ωsp = 0.685 rad/sec (with n/α = 3.9) is Level 2 on figure 13c; and τe = 0.186 is Level 2.
What appeared to be an adequate augmentation (figure 86) results in an airplane that is not much better
than the basic F-16. Note that the equivalent short period is lower than either ω1 or ω2 in figure 86 - an
illustration of the fact that it is incorrect to pick a dominant root to plot on the figure 13 boundaries.
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MIL-STD-1797A
APPENDIX A
FIGURE 86. AFTI/F-16 θ → Fs for IBU (q → δe closed).
In its application to direct force control modes, AFWAL-TR-81-3027, the bandwidth criterion was found to
work in areas where conventional criteria are inappropriate. This is discussed in more detail in 4.6.1.3.
Another lesson is that many pilots dislike having to push forward to get the aircraft on the ground at the
end of a landing flare, a characteristic noticed on a number of attitude-hold systems. They can learn to do
it, but it is against training and instinct.
Rynaski (Calspan Final Rpt No 7205-8) and Hoh (“STOL Handling Qualities Criteria for Precision
Landing”) point out that the location of the integrator inverse time constant with respect to 1/ Tθ2 [ figure
87], and the low-frequency residue, can alter the classical α response, in which a step control input soon
produces a constant angle-of-attack increment. The q feedback will tend to cancel some poles with zeros
of the q/qe transfer function to give a normal-looking mid-frequency response, and also to suppress the
low-frequency phugoid response; but the zeros of α/qe and γ/qe do not cancel any closed-loop poles. As a
result, a conventional-appearing θ/qcmd response can easily be obtained but the α/qcmd and γ/qcmd transfer
functions must still have an extra pole-zero pair. Only if 1/Tq is chosen to be near 1/ Tθ2 can the α
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response to a step command be kept from ramping off instead of reaching a steady value, as it does in
the conventional case. Unless 1/Tq is so chosen, there will no longer be a consistent relationship between
γ and θ (since θ) is still α + γ), so that the common piloting technique of controlling γ through θ will be
more difficult. This inconsistency could be important in tasks involving precise control of the flight path.
FIGURE 87. Frequency and time response comparison.
5.2.1.2 Short-term pitch response-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.1.2)
As the design progresses, analysis, simulation and flight test will be appropriate demonstration means. In
the end, flight tests will measure some parameters directly and collect data to validate analytical and
simulator results for the rest.
For Level 3. the time to double amplitude must be obtained using the perturbation values of pitch attitude
(or airspeed) away from trim. (A step input establishes a new equilibrium airspeed, but an impulse or
doublet does not. Thus for some seconds after a step control input the difference between stability and
instability may not be discernible.)
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APPENDIX A
VERIFICATION GUIDANCE
The related MIL-F-8785C requirements are paragraphs 3.2.2.1, 3.2.2.1.1 and 3.2.2.1.2.1.
a. Analysis
As a guide to critical flight conditions, here are some simple analytical approximations for some
longitudinal dynamics parameters. Values of the corresponding equivalent-system parameters, however,
may be quite different because of higher-order effects [This possibility applies to transfer-function zeros
(e.g., 1/Tθ2) as well as poles] even if there is only one pitch controller.
hn =
& Cmα/ CNα , % c /100
Static margin:
if Mu =
& 0
(More generally, hn is related to dδe/dV in stabilized straight flight)
Maneuver margin:
hm =
& -[Cmα/ CNα + Cmq, gρ c /(4W/S)], % c /100
Phugoid:
ω2p =
& 2(g/V)2 hn/hm + thrust & Mach effects, r2/s2
2ζpωp =
& 2(g/V)(CD/CL) + thrust & Mach effects, r/s
Short period:
ω 2sp =
& ghm/( CL1 k y c ), r2/s2
2
2ζspωsp =
& [CNα - (Cmq + Cmα)/(2ky)2]
n/α =
& CNα/CL1
g
, r/s
VCL1
1/Tθ2 =
& (g/V) (n/α)
rise time ∆t =
& 1/( Tθ2 ω 2sp )
where
CL1 =
k 2y = Iy/(m c 2)
W/( q S),
Frequency-response magnitude and phase
Factor
(jω+1/τ)±1=Me
jΦ
Exact
2
2
M=(ω +1/τ )±
Approximate
M =
& (1/τ)±1
1/2
Φ =
& ±τω
Φ = ± tan-1(τω)
for ω << 1/τ
}
M =
& ω±1
Φ =
& ± π/2 m 1/(τω)
[(jω)2+2ζωnjω + ω n ] ± =
MejΦ
2
M=[( ωn2 -ω2)2+4ζ2 ω n ω2]±1/2
2
Φ = ±tan-1{2ζ(ω/ωn)/[1-(ω/ωn)2]}
273
for ω >> 1/τ
}
M =
& ωn±2
Φ =
& ± 2ζ(ω/ωn)
for ω << ωn
}
MIL-STD-1797A
APPENDIX A
M =
& ω±2
Φ =
& ± π m 2ζωn/ω
274
for ω >> ωn
}
MIL-STD-1797A
APPENDIX A
The peak overshoot ratio for pitch response to a step control input can be estimated from figure 88. It is
apparent that overshoot are also possible with ζsp > 1 or ωsp Tθ2 < 1. A pure time delay, τ, Laplace
transformed as e-τs, gives a frequency-domain phase lag of φ = -τω and a constant magnitude of 1 ( 0 db).
FIGURE 88. Maximum pitch rate overshoot for a step control input (from DOT/FAA/CT-82/130-II).
In determining the validity of equivalent-system parameter estimates, the kinds of feedback and
compensation should be considered. In a simple case, matching only the pitch transfer function may be
adequate for 4.2.1.2. But generally, for the most accurate parameter identification all the data available
should be used. Pitch rate and normal acceleration only serve in more complex cases as a minimum to
assure kinematic relationships that match the assumptions.
McRuer and Myers (DOT/FAA/CT-82/130-II) show that for a superaugmented aircraft using high-gain
pitch-rate feedback with forward-loop integration (an equalization element Kq(s + 1/'Tq)/s] gives a normallooking short-period pitch transfer function except that its parameters are approximately
ωn2 =& KqMδ/Tq
ζ =
& Q K qTqMδ / 2
Numerator 1/T =& 1/Tq
Mδ = q S c Cmδ/Iy
crossover frequency (where M = 0 dB) of
high-gain asymptote:
ω c 0 =& k qM δ
Such augmentation tends to eliminate the phugoid mode, giving a very flat pitch response at lowfrequency (such an aircraft will be stable, although the u, α, γ and θ responses to control inputs will have
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APPENDIX A
a pole at s = 0). Care must be taken to account for all the contributions of flight control system
components to the equivalent time delay.
While normally the trim nz makes little difference in longitudinal dynamic response, there can be
exceptions as aerodynamic characteristics vary with speed or Mach number and angle of attack. Rollinduced inertial coupling is the subject of 4.8.1. Especially at low speed or high angle of attack, sideslip or
even coordinated turns can sometimes produce altered, coupled motions. The requirements apply
throughout the flight envelope, and for all amplitudes of motion. Both flight- and computer-generated time
histories, for example, can be Fast-Fourier-Transformed to get the equivalent linear-model parameters
that best represent the full nonlinear, time-varying motion.
In order to meet the Neal-Smith pilot-in-the-loop criteria, the designer must succeed in finding a
combination of Kp, Tp1, and Tp2 which will cause the amplitude and phase of YpYc to plot in the Level 1
region of figure 68. It is necessary therefore to perform a parameter search. This search procedure is not
difficult; it can be performed graphically using aids described in AFFDL-TR-70-74, or the process can be
mechanized on a digital computer. Because the calculations involved in evaluating the magnitude and
phase of YpYc as a function of frequency are simple to perform, it is feasible to use a simple trial-and-error
approach to test whether or not a proposed aircraft design meets the criteria.
Construction of frequency responses for matching or plotting is conveniently performed by linearizing the
high-order system (for all possible input amplitudes, if necessary). The linearized high-order model is
almost always available because it is used in the design process. If it is not and, for example, a flight
control element is to be changed on an existing system, and relinearization is not feasible, then fast
fourier analysis of a nonlinear simulation model of the system works well (as discussed below). The exact
linearized model must account for the lower-frequency effects of structural filters, aliasing, etc. in order to
faithfully represent the response in the frequency range of primary interest to flying qualities, past 10
rad/sec.
b. Simulation
Fast Fourier analysis of real-time or non-real-time simulations of the aircraft is best performed using
responses to a stick force input with wide frequency content. Background can be found, for example, in
“Applied Time Series Analysis, Vol. I, Basic Techniques."
c. Flight Test
Flight testing goals are to identify the aircraft, evaluate its operational merits, and determine specification
compliance. With limited flight time and test resources available, it is a challenge to get the data needed.
Only a very limited amount of compliance will be demonstrated directly. For the rest, flight validation/
correction of analytical models of the aircraft will make further checking possible.
AIAA Paper 80-1611-CP, AFFTC-TR-77-27, AIAA Paper 80-1633, and AFWAL-TR-81-3027 describe fast
Fourier reduction of flight data. AFFTC-TR-77-27 describes AFFTC experience with the method. AIAA
Paper 80-1633 discusses use of an electronically generated frequency sweep which worked adequately,
and AIAA Paper 80-1611-CP shows that FFT can work adequately even when the test condition is
theoretically least suited to the method. AFWAL-TR-81-3027 shows a pilot-generated frequency sweep
that worked very well. A typical frequency sweep and the resulting Bode plot (for a direct side force
control configuration from AFWAL-TR-81-3027) are shown on figure 89. The instrumentation required to
obtain these data was minimal, consisting of a yaw rate gyro and a pedal position transducer.
d. Piloting Aspects of Flight Test for Augmented Aircraft
AIAA Paper 80-1611-CP discusses the piloting aspects of flight test for augmented aircraft, from which
the following is extracted. The same factors apply to ground-based simulation.
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APPENDIX A
Figure [90] illustrates a landing time history of a configuration with 0.17 seconds actual
delay in the longitudinal command path. The landing is reasonably routine. Figure [91]
shows the same configuration, with the same pilot, on a different landing. A pilot-induced
oscillation, with virtual loss of control, is evident. As discussed by Smith [AFWAL-TR-813116], a high rate of descent had developed which forced the pilot to control the aircraft
more urgently…. The pilot awarded a rating of 5 presumably on the basis that the aircraft
had been landed routinely, with some deficiencies, on two occasions, and control was
almost lost on one landing due to one of those momentary aberrations which afflict pilots
for reasons unknown.
During the simulation the pilot in question proved himself to be adaptable to widely
different dynamics, whereas the main evaluation pilot in the same program, for example,
registered a more consistently progressive deterioration in rating as the dynamic flying
qualities parameters of the aircraft were degraded. The two pilots, though both highly
skilled, therefore demonstrated a contrast in piloting technique. This contrast is significant
because both adaptability and consistency are qualities which are needed, and therefore
commonly exhibited, by many development test pilots. The adaptive technique, however,
presents more of a challenge to the flying qualities engineer. He must pay particular
attention to pilot briefing and to choice of piloting task.
Pilot Briefing - Augmented dynamics possess potential problems which might not appear
unless the pilot adheres to the properly chosen demanding task. Therefore, the briefing
should encourage the pilot to tackle the task aggressively but realistically. If the pilot is
not aware of Smith's discussion of flying qualities cliffs, the briefing should include it
[AFFDL- TR-78-122]. The classic cliff example is the peculiarity of lags in augmented
dynamics, which can produce excellent flying qualities in loosely defined tasks, but pilotinduced oscillations in tightly-defined tasks. Therefore, the pilot should be encouraged to
demand much of the aircraft.
Piloting Task - A demanding but realistic task must be flown to expose potential flying
qualities problems. An offset precision touchdown has proved very suitable for exploring
longitudinal landing dynamics, for example. However, this is not necessarily the critical
task - for lateral dynamics. Task selection is difficult because pilot's perceptions of
difficulty are sometimes misleading: the approach is commonly considered more difficult
than flare and touchdown, for example, whereas the touchdown phase can clearly be
critical [AFFDL- TR-78-122] .... There is an obvious need for operational realism in tasks,
though there is some evidence that deliberately unrealistic tasks such as handling
qualities during tracking (HQDT), might conveniently predict... difficulties in other more
realistic tasks [AFFTC-TD-75-1].
The values of τp and ωBW required to demonstrate compliance with the figure 57 boundaries are obtained
from open-loop frequency responses of pitch attitude such as those shown in figures 15, 17, and 19.
These plots may be obtained from analyses (figure 59) or from Fourier-transformed flight test or Simulator
data such as was shown in figure 89. The Air Force Flight Test Center (AFFTC) has had considerable
success in Fourier transforming flight test data taken during operational tasks (as opposed to specially
tailored frequency sweeps). This saves flight test time and allows configuration identification at the flight
condition to be utilized operationally.
If significant nonlinearities are present in the system, the open-loop frequency response will depend on
the size of the input used in the identification process. When such nonlinearities are suspected, several
frequency sweeps should be accomplished with different input magnitudes. Data taken during operational
tasks will implicitly account for nonlinear effects if technically good data can be obtained.
277
MIL-STD-1797A
APPENDIX A
FIGURE 89a. Typical DFC control frequency sweep.
FIGURE 89b. Fourier transformed heading response.
278
MIL-STD-1797A
APPENDIX A
FIGURE 90. Approach and landing, no pilot-induced oscillation, configuration P12
of AFWAL-TR-81-3116, medium offset approach (75 ft lateral, 50 ft
vertical), landing no. 1 (from AIAA Paper 80-1611-CP).
279
MIL-STD-1797A
APPENDIX A
FIGURE 91. Pilot-induced oscillation at touchdown.
280
MIL-STD-1797A
APPENDIX A
VERIFICATION LESSONS LEARNED
Finding the instantaneous center of rotation for control inputs from time histories is usually not difficult, if
normal acceleration is recorded at two or three longitudinal locations. However, in AFTI/F-111 analyses
and simulations a combination of multiple flight control surfaces and aeroelasticity complicated the task.
But the exact location is not critical; the goal is just to get the two nz zeros well beyond the frequency
range of piloted control.
Both theoretically and in practice, equivalent-system parameters determined by matching θ/δes or θ/Fes
with fixed 1/Tθ2 should be about the same as those determined by simultaneously matching θ and nzCR
with free 1/Tθ2.
In practice a true step is not usually feasible. Time may be measured from the midpoint of the control
input transient, as sketched, for the most abrupt input feasible. Good instrumentation will be needed to
measure the time delay accurately.
281
MIL-STD-1797A
APPENDIX A
4.2.2 Pilot-induced pitch oscillations.
The pitch attitude response dynamics of the airframe plus control system shall not change abruptly with
the motion amplitudes of pitch, pitch rate or normal acceleration unless it can be shown that this will not
result in a pilot-induced oscillation. The total phase angle by which normal acceleration measured at the
pilot's location lags the pilot's pitch control force input at a criterion frequency, ωR, shall be less than
___________.
REQUIREMENT RATIONALE (4.2.2)
The purpose of this requirement is to insure that aggressive tracking behavior will not result in instabilities
of the closed-loop pilot/aircraft system. Any such tendency will degrade or even destroy mission
effectiveness and likely will be dangerous.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.2.2.3.
Recommended value: When a quantitative requirement is desired, the specified phase angle should be
less than 180 - 14.3 ωR degrees when ωR is in radians/second. The criterion frequency ωR is defined to be
any frequency within the range 1 < ωR < 10 rad/sec at which lightly damped (resonant) oscillations in pitch
attitude can result from turbulence inputs or from piloted control of the aircraft when used in the intended
operational manner. This requirement should be waived at the discretion of the procuring activity for those
flight conditions for which the ratio of normal acceleration measured at the pilot's location to pitch rate,
evaluated at the criterion frequency, is less than 0.012 g/deg/sec.
A related requirement is 4.2.8.2. Also, see 4.1.11.6 for a general PIO requirement. The qualitative
requirement of MIL-F-8785C is generalized in view of uncertainties in the state-of-the-art of flight control
system design, a tacit recognition of the complexity of the PIO problem; no detailed specification is, at this
time, a guarantee against building a PIO-prone airframe/flight-control-system combination.
The requirement precludes PIO, PIO tendencies or general handling qualities deficiencies resulting from
amplitude-dependent changes in aircraft dynamic response to pilot control inputs. These effects can be of
mechanical origin, e.g. bobweights coupled with static friction, or due to saturation of elements within the
control system, or due to compensation added to the automatic control system. PIO has occurred in the
T-38A, A-4D, and YF-12 due to such abrupt changes. Other known sources are short-period dynamics
(e.g. large ωsp Tθ2), feel system phasing (e.g. effective bobweight location not far enough forward), and
sensitive control force and motion gradients. AFFDL-TR-69-72 and Norair Rpt NOR-64-143 can furnish
some insight.
We are currently in somewhat of a quandary regarding a specific requirement for PIO. It would, in fact,
seem that the equivalent systems and bandwidth requirements (4.2.1.2) as well as the transient Fs/n
criterion (4.2.8.2) were specifically formulated to insure that piloted closed-loop tracking in the pitch axis
would be satisfactory. Hence, this requirement to some extent seems redundant. The following discussion
of the proposed criterion, originally presented in AFFDL-TR-77-57, is taken from AFWAL-TR-81-3109.
The PIO theory of AFFDL-TR-77-57 postulates that if the pitch (θ) loop is resonant at frequency ωR, then
the pilot may at some time (which cannot necessarily be predicted) attempt to control normal acceleration
azp to the exclusion or near exclusion of θ. According to AFFDL-TR-77-57, a PIO may occur when the
open-loop pilot - vehicle normal-acceleration response nz (jω)/nz e (jω) (the subscript e denotes the error
sensed by the pilot) is subjectively predictable: concentrated about some resonant frequency within the
pilot's bandwidth of control, with a magnitude there above a threshold value. This situation may arise
during pitch target tracking or as a result of the pitching response to a large, abrupt control input, failure
transient or gust. A pilot attempting to control normal acceleration at that frequency will incite a PIO if no
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MIL-STD-1797A
APPENDIX A
phase margin exists there; that is, if the phase angle of the nz (jω) transfer function is as negative as -180
deg at the resonant frequency. Using a pure 0.25 sec time delay plus gain to model the pilot, the stated
phase requirement for the aircraft is evolved. Violation of the phase criterion implies that if the pilot
switches to azp control, the acceleration loop will be dynamically unstable and a PIO will be initiated. This
paragraph provides the flight control system engineer with a quantitative criterion for minimum required
dynamic performance of feel and control systems.
The minimum amplitude cited is proposed as a quantitative guide for preliminary identification in the
design process (airframe or flight control system) of a threshold of pilot sensitivity, below which PIOs are
unlikely. A combined threshold is postulated of maximum acceptable rms pitch rate in tracking and
minimum az consciously felt by the pilot. More data should be collected from in-flight simulation to
establish the validity of this response ratio; the number selected, 0.012 g/deg/sec, conforms to past cases
of longitudinal PIO (AFFDL-TR-77-57).
The frequency ωR is, in disguise, a closed-loop, pilot/vehicle parameter. Fortunately it is also a very
physical parameter (pitch loop resonant frequency) that is readily understood and accepted. No method is
given in the standard for its selection; methods for doing so are contained in AFFDL-TR-77-57. The
frequency ωR can be readily identified from flight test.
The existence of a significant resonance in closed-loop pitch attitude control indicates that the pilot has
closed the loop with very little phase or gain margin. It is difficult to conceive how such closures would
occur on aircraft that meet the Level 1 equivalent system or bandwidth boundaries (4.2.1.2).
Gibson's views of PIOs are taken directly from ICAS-86-5.3.4:
High order characteristics are associated with pilot-vehicle closed loop handling problems
or PIO. As this term has been used to describe low order problems, the differences
should be clearly understood. The abrupt pitch bobble type is discontinuous, consisting of
repeated tracking corrections. The sluggish pitch overdriving type is also discontinuous
with input pulses to stop the unpredictable excess in response. Although the aircraft is
not under complete control, it is not out of control.
High-order PIO is a continuous out of control attitude instability, the amplitude ranging
from small to large and potentially destructive. Because the problem is due to inadequate
pilot-vehicle closed loop gain and phase margins, examination of the pitch attitude
frequency response identifies the cause and the solution.
Figure 76 shows the features which separate low and high-order pitch handling. The area
of interest can be confined to the region of phase lags between 180 and 200 degrees
which determines the PIO frequency. This arises from the success of the synchronous
pilot (NOR-64-143) in PIO analysis, assuming that any pre-PIO equalization is
abandoned for a pure gain behaviour in the undamped or divergent oscillation. The
correct frequency is adopted instantaneously with the stick in phase with the pitch attitude
error and 180 degrees out of phase with the attitude. The stick is not always moved so
purely in practice, but very often the pilot can be seen to apply the stick a little too quickly
and then hold it while waiting for the pitch rate reversal before also reversing the stick.
The tendency of a configuration to PIO can therefore be assessed without using a pilot
model by empirically establishing the range of characteristics found in actual PIO
examples. Enough have now been published to do this with considerable accuracy. An
important feature at the PIO frequency is the response gain. If this is small enough,
dangerous oscillation amplitudes cannot occur, and PIO has not been found where this is
less than 0.1 degrees per pound of stick force. This is not a completely necessary
condition but it is a highly desirable design aim.
283
MIL-STD-1797A
APPENDIX A
PIO's have occurred most frequently, though not exclusively, in the landing flare. The
connection with the commonplace stick pumping is well established. This subconscious
excitation of pitch acceleration in the flare occurs near the same frequency as a PIO. If
the attitude in the oscillation suddenly intrudes into the pilot's awareness, a ready-made
PIO is already in existence. The lower the frequency, the large is the attitude oscillation at
the usual acceleration amplitude of about 6 deg/sec2, and the more likely the conversion
becomes. This indicates strongly the desirability of a high crossover frequency through
the PIO region.
While an oscillation amplitude of less than 0.5 degree in the flare will not usually be
noticed, the one significantly more than a degree is very likely to, this or the
corresponding pumping/PIO frequency is not an ideal parameter for correlation. The most
successful has proved to be the rate at which the pitch attitude phase lag increases with
frequency in the PIO lag crossover region, equally applicable to the landing or to target
tracking tasks. By the nature of the attitude frequency response, if the crossover
frequency is low and the attitude attenuates only slowly towards the crossover region, the
phase rate is large. If the frequency is high and there is substantial attenuation, the phase
rate is low. The gain margin is increased, the stick pumping amplitude is reduced and the
tendency to PIO is decreased automatically by designing a low phase rate into the control
laws.
This simple attitude parameter alone is almost sufficient to quantify the tendency to high
order PIO, and it correlates well with available examples of high order PIO. Figure 77
shows the trends, with an accuracy good enough to allow Level 1, 2 and 3 boundaries to
be drawn, if desired. For the control law designer it is enough to aim for a phase rate of
less than 100 degrees per cps and attitude response phase rate of less than 100 degrees
per cps and attitude response smaller than 0.1 deg/lb at the crossover. These
characteristics are a natural feature of low order aircraft whose attitude phase lag
exceeds ISO degrees due to the power control and so could in principle suffer from PIO,
yet do not. Early examples of bobweight PIO were high-order in kind and are found to
have had very large phase rates with the stick free.
For most combat aircraft configuration, consideration of normal acceleration effects does
not improve the PIO analysis. The g at the cockpit is usually attenuated and phase
advanced relative to the cg and will often not reach the 180 degrees lag necessary for
piloted instability. Human sensing of the g response is poor and at the initiation of the PIO
the g may be undetectable. In large aircraft with the cockpit far ahead of the cg, the
heave can have a significant effect and has to be taken into account in the dominant
requirement to optimize the pitch attitude behaviour.
Although the attitude to stick force response gain is significant in PIO, there is little
evidence that a damper modifies the pilot's stick phasing in a PIO and only the stiffness
component should be used. Where PIO tendencies exist, they will be exacerbated by a
high stick stiffness. Gradients of 5 to 8 lb/in with forces of 2 to 2.5 lb/g have proved to be
extremely satisfactory for [fly-by-wire] aircraft. Designed to the phase rate and gain
margin criteria discussed above, the attitude gain phase rate and gain at the PIO
frequency is only some 0.5 deg/in. In AFFDL-TR-74-9 case 4D had high phase rate and
low PIO gain margin. With a gradient of 22 lb/in and 6.7 lb/g it had an attitude gain of 7
deg/in at the PIO frequency. Not surprisingly it suffered from continuous pitch oscillations
and severe tracking PIO, earning ratings of 9 and 10.
284
MIL-STD-1797A
APPENDIX A
The boundaries in the frequency response criteria of figure [75] are based directly on
these considerations and will eliminate high order PIO. Low order PIO will also be
eliminated by the optimization criteria given above.
SUPPORTING DATA
AFFDL-TR-77-57 illustrated several examples of PIO-prone aircraft. One example is similar to the YF-17
as simulated on the USAF/Calspan variable stability T-33. The θ and azp transfer functions are given as
follows4;
K θ (2) (2.3) [.44,11.]
θ
=
Fs (0) (5) [.89,1.98] [.7,4.0]
a zp
Fs
=
K a (2) (2.3) [.08,5.04] [.44,11.]
(.9) (5.) [.89,1.98] [.7,4]
The following discussion is quoted from AFFDL-TR-77-57.
Figure 92 is a Bode plot of the airplane's pitch attitude dynamics, θ/Fs(jω). Assuming that the crossover
frequency will lie between 2 and 4 rad/sec, it is clear that the aircraft dynamics are roughly of the form
K/s2 in this region. As a rule, dynamics of this sort will lead to lightly damped closed-loop oscillations and
degraded pilot opinion ratings. An inspection of the data base of AFFDL-TR-65-15 and a modicum of
iteration suggest that a reasonable model for pilot dynamics in pitch tracking would be
Yp (jω) = Kp (2.5jω + 1) e-.385jω
A Bode pilot of the open-loop system dynamics YOL = Yp (θ/FS) is also shown on figure 92. Figure 92
indicates that the absolute maximum crossover frequency with the Yp(jω) is 3.3 rad/sec. Accordingly, ωc =
2.9 was selected and is assumed to be consistent with what would be measured in actual flight; this,
yields a small phase margin (about 16 deg). Obviously, even small increases in pilot gain will rapidly
degrade system stability. This result appears to be consistent with the evaluation pilots' comments about
the poor pitch handling qualities of this configuration in flight tests (AIAA Paper 75-985).
4
(1/T) → [s + 1/T]; [ζ, ω] → [s2 + 2ζωs + ω2]
285
MIL-STD-1797A
APPENDIX A
The corresponding closed-loop dynamics, θ/θc = YCL, are shown on figure 92 for ωc = 2.9. Obviously the
closed-loop system is extremely resonant at this condition. It is evident by inspection that the resonant
peak of θ/θc will dominate the azp power spectrum. The corresponding damping ratio for this mode is
approximately 0.03. Thus, by the simplified criterion for subjective predictability, it must be concluded that
PIO cannot be ruled out on the basis of pitch control handling qualities. The resonance frequency ωR is
3.0 rad/sec for the given Yp(jω) More pilot lead and higher pin would increase ωR somewhat.
FIGURE 92. YF-17 pitch attitude dynamics (AFFDL-TR-77-57).
By the assessment rules of AFFDL-TR-77-57, the analysis must now proceed to an investigation of
stability of the azp → Fs loop when the pilot's gain is adjusted to make ωc = ωR. The total azp → Fs system
phase (φ) versus frequency is plotted on figure 93 in accordance with the rules of the PIO theory. The
pilot time delay was assumed to be 0.25 seconds. At ω = 3.0 we have φ = -205 deg, 180 + φ = -25 deg
(the system phase margin), and we see that the acceleration-closed loop is unstable. Thus, longitudinal
PIO can be initiated provided that the pilot attempts to control azp.
286
MIL-STD-1797A
APPENDIX A
FIGURE 93. YF-17 acceleration control system dynamics (AFFDL-TR-77-57).
& | (3.0j) is 0.031 g/deg/sec. Thus by present theory we would be justified in concluding
The ratio |azp / θ
PIO would be likely with this airplane and control system.
& | (3.0j) = 0.0213
The actual normal acceleration dynamics simulated with the NT-33A yield |azp / θ
g/deg/sec. This is about twice the criterion value of 0.012; on that basis it can be concluded that errors in
the simulation of azp motion amplitude were probably of no consequence.
The PIO frequency and amplitude obtained with the NT-33A simulation are unpublished. It is known from
informal communication between the writer of AFFDL-TR-77-57 and Calspan staff members that the PIO
frequency was approximately 1/2 cps. It may therefore be concluded that this analysis (and, as a
consequence, the present theory) is supported by the flight test results.
REQUIREMENT LESSONS LEARNED
The example given in the Supporting Data showed that the criterion successfully predicted a PIO. But,
what if we checked the pitch dynamics against the equivalent systems or bandwidth criteria of 4.2.1.2? A
lower order equivalent system was not run for the dynamics presented. However, defining the shortperiod damping as 0.89 (as is done in AFFDL-TR-77-57) may not be appropriate considering the
significant number of higher-order modes that exist. The bandwidth criterion can be checked directly from
figure 92 (θ/Fs) with the following results:
287
MIL-STD-1797A
APPENDIX A
ωBW = 1.3 rad/sec
τp = 0.16 sec
Plotting these values on figure 56 shows that the aircraft is very close to the Level 3 region of the flying
qualities boundary. Hence the conclusion that the aircraft is PTO prone is not surprising. In fact, the
resonant peak in YCL of figure 92 is a direct consequence of these particular values of ωBW and τp.
Nonetheless, it may be desirable to retain the criteria to emphasize the notion that azp may well be a key
parameter for identifying PIO-prone aircraft. Also, it may be possible for a configuration to pass the lowerorder equivalent system or bandwidth criterion and be caught by the PIO criterion.
A very good summary report on PIOs is given in NOR-64-143. The following paragraphs from that
reference discuss the causes of PIOS:
There are several ways of looking at the causes of a PIO. One is to catalog all the PIO
situations ever recorded, including all the necessary subsystem details, etc., and then to
say that each combination of vehicle and subsystem when combined with the pilot was
the cause of a PIO. Another way is to note that certain system phenomena such as stickforce-to-control-deflection hysteresis often lead to PIO when other conditions are right. A
third way, and one which seems to transcend the difficulties of the previous two, is to say
that certain inherent human physical limitations are the basic cause for any PIO. This is
not to degrade the human pilot's role but, instead, to emphasize it, because it is unlikely
that any black-box could be devised which is as clever and effective in coping with
unmanageable controlled elements as a skilled pilot. Were it not for the pilot's versatile
gain adaptability, many flight conditions would be unstable. But there is a limit to the
rapidity with which the human can adapt, and this can sometimes lead to a PIO.
When referred to the pilot, then, the basic causes of PIO seem to fall into the following
categories:
1. Incomplete pilot equalization
a. Incomplete training
b. Inappropriate transfer of adaptation (i.e., carry over of improper techniques from
another aircraft)
2. Excessive demands on pilot adaptation
a. Required gain, lead, or lag lie outside the range of normal capabilities
b. Rate of adaptation is too slow to preclude oscillation
c. Inadequate capability to cope with system nonlinearities
3. Limb-manipulator coupling
a. Impedance of neuromuscular system (including limb) on control stick or pedals
changes feel system dynamics
b. Motion-induced limb force feedback (e.g., arm becomes a bobweight)
Table XIV, from NOR-64-143, lists some known PIO cases and their causes for then-current (early 1960s)
aircraft. The causes are equally relevant for modern aircraft, and the lessons learned from the cases
listed are valuable in preventing PIOs.
288
MIL-STD-1797A
APPENDIX A
5.2.2 Pilot-induced pitch oscillations-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.2)
It would be an easy matter for SPO engineers to ascertain compliance with this paragraph without relying
on pilot/vehicle analysis methods. For example, ωR and the specified phase lag can easily be obtained
from simulator or in-flight time histories (ground-based simulations will not show up accelerationdependent PIO tendencies). Nonetheless, analytical estimates can -- and should - be made by the
airframe manufacturer as part of the design evolution. For flight evaluation, the PIO rating scale of figure
12 will be helpful.
VERIFICATION GUIDANCE
The user should refer to Chapters IV, V, and VI of AFFDL-TR-77-57 when applying the quantitative
requirement. PIOs are associated with precise tracking as in air-to-air gunnery, formation flying, flare, and
touchdown. PIOs observed in flight are often not obtained in ground-based simulators, even ones with
some motion. Tight, aggressive pilot control action will tend to bring on any PIO tendencies. High
sensitivity to control inputs is often a factor. Some pilots are more PIO-prone than others, depending upon
piloting technique.
VERIFICATION LESSONS LEARNED
These requirements are an attempt to catch and correct any PIO tendencies as early as possible in the
design, when changes are easiest and least costly to make. They also have been found helpful in
identifying PIO tendencies in flight and determining fixes.
289
MIL-STD-1797A
APPENDIX A
TABLE XIV. Classification of some known PIO cases (from NOR-64-143).
Examples shown as: SPECIES (Aircraft); Critical Subsystem*; Critical Flight Condition**: Remarks
TYPE
CLASS
I. LINEAR
II. SERIES NONLINEAR ELEMENTS
PITCH
III. SUBSIDIARY FEEDBACK NONLINEAR
ELEMENTS
IMPROMPER SIMULATION; D, V; a:
Abnormally high value of 1/Tθ2 and low (ζω)sp
led to zero ζsp when regulating large
disturbances.
PORPOISING (SB2C-1); F; c: Hysteresis in
stick versus elevator deflection resulted in lowfrequency speed and climb oscillations.
BOBWEIGHT BREAKOUT (A4D-1, T-38A): F,
B; a: At high-g maneuvers the bobweight
overcomes system friction and reduces
apparent damping of the aircraft in response to
force inputs, resulting in large oscillations at
short-period.
GCA-INDUCED PHUGOID (C-97); D; c, b:
Lag from radar-detected error to voice
command led to unstable closed-loop
phugoid mode.
J. C. MANEUVER (F-86D, F-100C); F, S; a:
Valve friction plus compliant cabling resulted in
large oscillations at short period.
LOSS OF PITCH DAMPER
ARM ON STICK (A4D-1, T-38A); F; a: Arm
mass increases feel system inertia, leads via
B feedback to unstable coupling with shortperiod dynamics if pilot merely hangs loosely
onto stick after a large input.
PITCH-UP (XF-104, F-101B, F-102A); V; c:
Unstable kink in M(α) curve led to moderateperiod oscillations of varying amplitudes
(depending on extent and nature of the kink)
during maneuvers near the critical angle-ofattack.
LANDING PIO (X-15); S; b: Closed loop around
elevator rate limiting caused moderate
oscillations at short-period.
LATERALDIRECTIONAL
ωφ/ωd EFFECT (X-15, NT-33A, F-101B, F106A, KC-135A, B-58); V; c: Zeros of
roll/aileron transfer function are higher than
Dutch roll frequency, | ωφ/ωd| > 1.0, leading
to closed-loop instability at low ζ d conditions.
LOSS OF YAW DAMPER
BORESIGHT OSCILLATIONS (F-5A); D, V;
c: Spiral roll mode driven unstable if roll
information is degraded during gunnery.
YAW
FUEL SLOSH SNAKING (KC-135A, T-37A);
V; c: Fuel slosh mode couples with Dutch
roll mode when rudder used to stop yaw
oscillation.
TRANSONIC SNAKING (A3D); V, F; a,c:
Separation over rudder causes control reversal
for small deflections, leading to limit cycle if
rudder used to damp yaw oscillations.
ROLL
NONE KNOWN
PILOT-INDUCED CHATTER (F-104B); A; c:
Small limit cycle due to damper aggravated
whenever pilot attempted to control it.
* Critical Subsystems:
D = Display
F = Feel system (except
B)
B = Bobweight
** Critical Flight Conditions:
a = Low altitude, near-sonic Mach
b = Landing approach and takeoff
S = Power servo actuator
V = Vehicle (airframe)
A = Augmentor (damper)
c = Cruise
291
MIL-STD-1797A
APPENDIX A
4.2.3 Residual pitch oscillations.
In calm air, any sustained residual oscillations shall not interfere with the pilot's ability to perform the tasks
required in service use of the aircraft. For Levels 1 and 2, oscillations in normal acceleration at the pilot
station greater than ____________ will be considered excessive for any Flight Phase. These
requirements shall apply with the pitch control fixed and with it free.
REQUIREMENT RATIONALE (4.2.3)
The requirement prohibits limit cycles in the control system or structural oscillations that might
compromise tactical effectiveness, cause pilot discomfort, etc. This requirement may be considered a
relaxation of the requirement in 4.2.1 for positive damping at all magnitudes of oscillation. Its intent is to
recognize thresholds below which damping is immaterial.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is paragraph 3.2.2.1.3.
The recommended value is 0.02g. Given the proper data, this threshold could be made a function of
frequency in order to correspond more closely with human perception.
REQUIREMENT LESSONS LEARNED
Allowable normal acceleration oscillations have been decreased to 0.02 g from the 0.05 g of MIL-F8785C. This is based on flight test experience with the B-1 (AFFTC-TR-79-2), which encountered limit
cycle oscillations during aerial refueling, subsonic and supersonic cruise. A primary contributor was
identified to be mechanical hysteresis in the pitch system. According to AFFTC-TR-79-2, “Flying qualities
were initially undesirable due to this limit cycle." Normal acceleration transients in cruise were about 0.05
- 0.12 g, as figure 94 shows. The limit cycle was eliminated by installation of a mechanical shaker (dither)
vibrating at 20 Hz.
5.2.3 Residual pitch oscillations-verification.
Verification shall be by analysis, simulation and flight test
VERIFICATION RATIONALE (5.2.3)
Limit cycle amplitude depends on characteristics of the actual hardware and software, and so may be
different in simulations than in actual flight. Measurements of normal acceleration at the pilot's station
should be made in the course of test flight to meet the other flying quality requirements.
VERIFICATION GUIDANCE
Residual oscillations are limit cycles resulting from nonlinearities such as friction and poor resolution.
Negative static stability will contribute and low damping may augment the amplitude. Thus high speed,
high dynamic pressure or high altitude may be critical. Residual oscillations are most bothersome in
precision tasks.
VERIFICATION LESSONS LEARNED
292
MIL-STD-1797A
APPENDIX A
FIGURE 94. Effect of dither on B-1 limit cycle oscillations (from AFFTC-TR-79-2).
293
MIL-STD-1797A
APPENDIX A
4.2.4 Normal acceleration at pilot station.
Normal acceleration at the pilot station due to pitch control inputs shall have the following characteristics:
_____________.
REQUIREMENT RATIONALE (4.2.4)
The dynamic normal acceleration response at the pilot station to the pitch controller should not be
objectionably large or of a confusing nature in terms of the pilot's perception of pitch rate response to a
pitch controller input.
REQUIREMENT GUIDANCE
This is a new requirement whose need is apparent, though insufficient information exists to formulate
recommended criteria. Unusual pilot locations can adversely affect handling qualities. A prominent
example is the Space Shuttle, in which confusing acceleration cues played a part in pilot-induced
oscillations encountered during approach and landing tests (NASA-CR-163108). In that vehicle the pilot's
station is noticeably aft of the instantaneous center of rotation for elevator inputs, so when he pulls back
on the controller he first sinks noticeably before rising.
In most cases, the pilot station is forward of the instantaneous center of rotation for pilot pitch control
inputs, giving an initial impulse 1x ⋅ q& o in the correct direction and a transfer function (short-term
approximation)
a zp
Fs
=
K[s 2 + 2ζ a ω a s + ω a2 ]
[s 2 + 2ζ sp ω sp s + ω 2sp ]
Generally ωa > ωsp, but it is conceivable this may not hold universally. Figure 95 shows no ill effects from
ωa = 1.25, ωsp < 1 for simulated approach and landing of a very large transport. The initial az seems to
help overcome -- to an extent -- any adverse time-delay effects, but a limit might be reached where the
initial response is too abrupt.
Blended direct-lift control has not always produced the expected gains. The F-16 CCV and AFTI
maneuver enhancement modes were well liked for both their quickened path response and their gust
alleviation. But evaluating washed-out spoiler blended with elevator control, Hanke, Wilhelm and Lange
(AGARD-CP-333) found deteriorating ratings with increasing DLC gain. Evidently, if the pilot must close
an inner attitude loop, he needs some phase separation between pitch and path responses in order to
distinguish the two during the approach phase of landing. For the classical case, this separation is given
by tan-lωsp Tθ2. Another in-flight evaluation, of approaches and actual landings with a rate
command/attitude hold system, failed to realize the improved ratings that DLC had produced in groundbased simulations--it has been argued whether the disturbing heave motion associated with stick,
motions was attributable to the basic DLC or was an artifact of the simulation (Mooij and Van Gool,
AGARD-CP-333)
REQUIREMENT LESSONS LEARNED
294
MIL-STD-1797A
APPENDIX A
FIGURE 95. Pilot rating vs pilot position - center of rotation (from AFWAL-TR-81-3118).
5.2.4 Normal acceleration at pilot station-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.4)
Any difficulty will be apparent in tasks requiring very tight control of flight path.
VERIFICATION GUIDANCE
The effective control arm is given by lc / c = Cmδ / CLδ, positive forward from the c.g. (negative for an aft
tail).
The instantaneous center of rotation for control inputs is at lCR = Zδ/(Mδ + Zδ M w& ) =
& Zδ/Mδ, positive
forward of the c.g. [Zδ = - q S CLδ /m, Mδ = q S c Cmδ /IY].
For the two-degree-of-freedom approximation (assuming constant speed) the normal-acceleration
transfer-function numerator for control inputs is
295
MIL-STD-1797A
APPENDIX A
a
Nδ z = − g
C Lδ 
l
1 − x

l CR
C L1 







g/ V
s 2 −

2k 2y C L1




C mα  

2l x l c 
c C mα 
C

1− c

+
−
−
C
1
C
m
m
N
&
α
α 
q

2 

l
C


l
C
C
c
c
N
l
g
N
c
N
α 

α 
α 
c

s −


1 − l x / l CR
k 2y c 2 C L1  1 − l x / l CR 










296
MIL-STD-1797A
APPENDIX A
at a distance lx forward of the c.g., where ky
response.
= Iy /(m c 2). Fuselage bending might also affect the
VERIFICATION LESSONS LEARNED
With the USAF-Calspan TIFS (Total In-Flight Simulator) airplane, attempts to control the value of 1/Tθ2 via
direct lift control have excited a structural mode which was very evident in the evaluation cockpit at high
DLC gain. In one recent case the result was a long time delay followed by a very abrupt response at the
cockpit, which the pilot rated unacceptable or uncontrollable. Since all aircraft have some flexibility, the
lesson is to be alert for aeroservoelastic effects on flying qualities as well as on the structure and the flight
control system.
297
MIL-STD-1797A
APPENDIX A
4.2.5 Pitch trim changes.
The pitch trim changes caused by operation of other control devices shall not be so large that a peak
pitch control force in excess of 10 pounds for center-stick controllers or 20 pounds for wheel controllers is
required when such configuration changes are made in flight under conditions representative of
operational procedure. Generally, the conditions of table IV will suffice for determination of compliance
with this requirement. With the aircraft trimmed for each specified initial condition, and no retrimming, the
peak force required to maintain the specified configuration change shall not exceed the stated value for a
time interval of at least 5 seconds following the completion of the pilot action initiating the configuration
change. The magnitude and rate of trim change subsequent to this time period shall be easily trimmable
by use of the normal trimming devices. These requirements define Level 1. For Levels 2 and 3, the
allowable forces are increased by 50 percent.
REQUIREMENT RATIONALE (4.2.5)
These frequently encountered pitch trim changes, if too large, can add to pilot workload at critical times
during a mission.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.6.3.1.
Table XV gives the recommended conditions (For aircraft with variable-sweep wings, additional
requirements should be imposed consistent with operational employment of the vehicle. Thrust reversing
and other special features also need to be considered). These are the trim changes that, when larger
than the limits specified, have been bothersome in the past. Crossfeeds and feedbacks in the stability,
and control augmentation system generally will reduce the magnitude of these trim changes. Wing
downwash and vertical placement of the engines are two of the determining factors. For thrust reversing,
configuration-dependent aerodynamics play an important role.
4.1.13 gives additional general trim requirements.
REQUIREMENT LESSONS LEARNED
The direction of the trim change can also be important, producing either helpful or unfavorable coupling.
In any case the magnitude should not be excessive.
5.2.5 Pitch trim changes-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.5)
The evaluation should be made in the manner expected in operational practice, rather than necessarily
holding everything else constant.
VERIFICATION GUIDANCE
Initial trim conditions are listed in table XV.
VERIFICATION LESSONS LEARNED
298
MIL-STD-1797A
APPENDIX A
1
Flight
Phase
Attitude
Approach
homin
2
3
TABLE XV. Pitch trim change conditions.
Initial Trim Conditions
Speed
Landing High-lift
Thrust
Gear
Devices
& Wing
Flaps
Normal
Up
Up
TLF
pattern
entry
speed
Up
Up
TLF
Down
Up
TLF
4
Down
Up
TLF
Vomin
Down
Down
Down
Down
TLF
TLF
Vomin
Down
Down
Down
Down
TLF
TLF
Down
Take-off
Minimum
flap-retract
speed
Up
Take-off
Take-off
thrust
Take-off
thrust
Speed for
level flight
Up
Up
MRT
12
Up
Up
MRT
13
Up
Up
MRT
Up
Up
TLF
5
6
7
8
Approach
9
Takeoff
10
11
Cruise
and airto-air
combat
14
homin
and
homax
Speed for
best range
Configuration
Change
Parameter
to be held
constant
Gear down
Altitude
and
airspeed*
Gear down
Extend highlift devices
and wing
flaps
Extend highlift devices
and wing
flaps
Idle thrust
Extend
approach
drag device
Takeoff thrust
Takeoff thrust
plus normal
cleanup for
wave-off (goaround)
Gear up
Altitude
Altitude
and
airspeed*
Retract highlift devices
and wing
flaps
Idle thrust
Altitude
Airspeed
Airspeed
Airspeed
Airspeed
Pitch
attitude
Airspeed
Pitch
attitude
Actuate
deceleration
devices
Maximum
augmented
thrust
Actuate
deceleration
device
*Throttle setting may be changed during the maneuver.
Notes: -
Auxillary drag devices are initially retracted, and all details of configuration not
specifically mentioned are normal for the Flight Phase.
If power reduction is permitted in meeting the deceleration requirements established for
the mission, actuation of the deceleration device in #12 and #14 shall be accompanied
by the allowable power reduction.
299
MIL-STD-1797A
APPENDIX A
4.2.6. Pitch axis response to other inputs.
4.2.6.1 Pitch axis response to failures, controls free.
With controls free, the aircraft motions due to partial or complete failure of any subsystem of the aircraft
shall not exceed the following limits for at least ________ seconds following the failure: ______________.
REQUIREMENT RATIONALE (4.2.6.1)
Quantitative limits are needed to avoid pilot workload increases and flight safety problems.
REQUIREMENT GUIDANCE
The applicable MIL-F-8785C paragraph is 3.5.5.1.
Recommended limits on transient motions within the first 2 seconds following failure are as follows:
Levels 1 and 2 (after failure): ±0.5 g incremental normal acceleration at the pilot's station,
except that neither stall angle of attack nor structural limits shall be exceeded. In addition,
for Category A, vertical excursions of 5 feet.
Level 3 (after failure): No dangerous attitude or structural limit is reached, and no
dangerous alteration of the flight path results from which recovery is impossible.
These limits were taken from paragraph 3.5.5.1 of MIL-F-8785C. Although the intent of the requirement is
to insure that dangerous flying qualities never result, there may be some benefit to a noticeable transient
after a failure, or after transfer to an alternate control mode, in order to alert the pilot to the change That
possibility is left to the designer without explicit direction to minimize transients. This requirement also
places quantitative limits on the altitude change, effectively restricting the 2-second average acceleration
in addition to the peak value.
REQUIREMENT LESSONS LEARNED
The revision to MIL-F-8785C followed the recommendations of Systems Technology Inc. TR-189-1: the
authors noted that the allowable transient levels of MIL-F-8785B were consistent with failure probability
considerations, but not with flying qualities considerations. Level 2 had a lower probability of occurrence
than Level 1 and was permitted to have larger transient responses; however, Level 2 is a poorer handling
qualities state and cannot accept the larger responses as readily. It was felt that the values in MIL-F8785C were representative of transients which could be handled with Level 1 flying qualities. Conversely,
the low allowable transients of MIL-F-8785B were conducive to soft failures which could lead to
catastrophic situations if undetected by the pilot. This comment applied to the B-58, in particular, and led
General Dynamics/Ft. Worth to suggest a minimum allowable transient (according to Systems
Technology Inc TR-189-1). This has not been incorporated into this document, but should be a
consideration in the design process.
5.2.6.1 Pitch axis response to failures, control free-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.6.1)
Some failures may be considered too dangerous to flight test. Where final demonstration is by simulation,
flight-validated aerodynamic data should be used with actual flight hardware and software, loaded as
necessary to replicate the response in flight.
300
MIL-STD-1797A
APPENDIX A
VERIFICATION GUIDANCE
Worst-cast flight conditions should be identified and tested. High control effectiveness, authority and gain;
low or negative static stability or damping; low weight and inertia will tend to make the transients larger.
Generally a dynamic analysis is needed, but constant speed can be assumed for the two-second period
of time.
Mδ = q S c Cmδ / Iy = Cmδ g c /(CL1 K 2y )
where CL1 = W/qS, Ky is the radius of gyration (dimensional) in pitch.
VERIFICATION LESSONS LEARNED
301
MIL-STD-1797A
APPENDIX A
4.2.6.2 Pitch axis response to configuration or control mode change.
The transient motions and trim changes resulting from configuration changes or the intentional
engagement or disengagement of any portion of the primary flight control system in equilibrium flight due
to pilot action shall be such that dangerous flying qualities never result. With controls free, the motion
transients resulting from these situations shall not exceed the following limits for at least seconds
following the transfer: These requirements apply only for Aircraft Normal States (4.1.6).
REQUIREMENT RATIONALE (4.2.6.2)
Pitch transients due to intentional mode switching are distracting. If the transients are too large, pilots
object.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.5.6 and 3.5.6.1.
Recommended values for transient motion limits within the first 2 seconds following transfer:
Within the Operational Flight Envelope: ±0.05 g normal acceleration at the pilot's station
Within the Service Flight Envelope: ±0.05 g at the pilot's station.
Since the intent of a flight control system is to improve the aircraft response characteristics--whether
measured by improved flying qualities or by increased mission effectiveness--any system which can be
chosen by the pilot should not cause noticeable transient motions. There has been some speculation as
to whether a small transient motion is or is not desirable. The argument for an intentional transient is that
inadvertent pilot switching of autopilot modes is less likely if accompanied by a noticeable transient
motion.
MIL-F-8785B allowed 0.05 g normal acceleration. This was increased to 0.10 g in MIL-F-8785C, in order
to allow if not encourage designers to provide some noticeable transient (see AFWAL-TR-81-3109). In
AFWAL-TR-81-3109 an accident was cited wherein the pilot inadvertently bumped off the altitude hold
mode (which automatically disengaged when a small force was applied to the control column). The flight
recorder showed a 0.04 g transient which went unnoticed by the crew, who were deeply involved in trying
to lower a malfunctioning landing gear. However, it is our condition that the undesirable features of
transient motions due to mode switching are significant. Furthermore, a distracted crew would probably
not notice a transient considerably larger than 0.04 g, especially if there were any turbulence at all.
Therefore, we are recommending that the maximum allowable transient of 0.05 g used in MIL-F-8785B be
utilized as in MIL-F-8785B and C; two seconds is deemed a reasonable time to allow for the pilot to
resume control.
REQUIREMENT LESSONS LEARNED
5.2.6.2 Pitch axis response to configuration or control mode change -verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.6.2)
Configuration changes and control mode changes are to be made by normal means.
302
MIL-STD-1797A
APPENDIX A
VERIFICATION GUIDANCE
Compliance should be evaluated at likely conditions for mode switching and at the most critical flight
conditions. Critical conditions will usually be the corners of the expected operational envelopes (e.g., a
SAS for power approach should be switched at the highest and lowest expected airspeeds, at low
altitudes). Some factors which determine critical conditions are given in the discussion of 4.2.6.1. Limited
analytical and ground-based simulation may be used to supplement actual flight testing, especially in the
early stages of development, but flight testing is ultimately required.
VERIFICATION LESSONS LEARNED
303
MIL-STD-1797A
APPENDIX A
4.2.7 Pitch axis control power
4.2.7.1 Pitch axis control power in unaccelerated flight.
In steady 1-g flight at all Service altituides, the attainment and holding of all speeds between VS and Vmax
shall not be limited by the effectiveness of the pitch control.
REQUIREMENT RATIONALE (4.2.7.1)
This requirement is intended to insure that the pilot can maintain equilibrium level flight throughout the
flight envelope by normal means.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.2.3.1.
Controllability at speeds down to the 1-g stall speed is generally deemed necessary for safety, as well as
full utilization, of maneuvering aircraft such as the military use. Vmax, the high-speed boundary of the
Service Flight Envelope, must be at least V0max; beyond that, it may be set by the contractor, who then
must deliver on his promise.
REQUIREMENT LESSONS LEARNED
5.2.7.1 Pitch axis control power in unaccelerated flight-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.7.1)
Operational flight test will help reveal any deficiencies in pitch control power. Compliance may be
demonstrated during measurement of the gradient of pitch control force with airspeed (4.2.1). The
controls are to be used in their normal manner, and sideslip minimized (its effect is specified separately).
VERIFICATION GUIDANCE
It is important to explore all corners of the V - h Service Flight Envelope. For example, a transonic tuck or
high-speed dives can be critical due to combined aeroelastic and Mach number effects. Extremes of
static stability or instability (Mach number, angle of attack, center of gravity) will be critical. Also, hinge
moments may limit control deflection and aeroelastic deformations may affect controllability. In 1-g
equilibrium level flight, net forces are zero and
Cm (α, δ, M, q , c.g.) + (zT / c ) [T/( q S)] = 0
VERIFICATION LESSONS LEARNED
304
MIL-STD-1797A
APPENDIX A
4.2.7.2 Pitch axis control power in maneuvering flight.
Within the Operational Flight Envelope it shall be possible to develop, by use of the pitch control alone,
the following range of load factors: _________. This maneuvering capability is required at constant
altitude at the 1-g trim speed and, with trim and throttle settings not changed by the crew, over a range
about the trim speed the lesser of 15 percent or 50 kt equivalent airspeed (except where limited by the
boundaries of the Operational Flight Envelope).
REQUIREMENT RATIONALE (4.2.7.2)
The pitch axis controller must be sufficiently powerful to produce an adequate range of load factors for
maneuvering. Fixed-wing aircraft generally use the pitch controller to affect flight-path changes.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is 3.2.3.2.
Recommended range of load factors:
Levels 1 and 2:
no (-) to no (+)
Level 3:
n = 0.5 g to the lower of:
a) no (+)
b) n = 2.0 for no(+) ≤ 3g
0.5 [no (+) + 1] for no (+) > 3g
The Level 1 and 2 values stem from mission performance needs; while the Level 3 values, which vary to
an extent with maneuverability of the aircraft, are related to flight safety.
The requirements for control effectiveness over a +15 percent range about the trim speed assure that
excessive amounts of pitch-surface-fixed static stability or instability will not limit maneuver capability
unduly, for any possible mechanization of the trim system. Where pitch control authority limit normalacceleration capability, the requirement at off-trim speeds often will be the designing consideration for
pitch control effectiveness. At most flight conditions, however, it would be expected that investigation at
the nominal V would be sufficient.
This requirement is restricted in application to the Operational Flight Envelope with relaxed requirements
for infrequent Failure States. Outside the Operational Flight Envelope, whatever falls out of the design is
acceptable, as long as the other control requirements are met (in particular, note the dive pullout and
control margin requirements). High supersonic speed and aeroelasticity and large high hinge moments at
high dynamic pressure tend to restrict the control capability. The Level 3 requirement assures modest
nose-down and nose-up control capability for stabilization as well as for altering equilibrium and
maneuvering. AFWAL-TR-83-3015 recommends requiring +1.5g capability for Level 3 at the design dive
speed, against the most adverse stabilizer (pitch) trim. That is consistent with some civil aircraft
requirements for a 2.5g transport: 0.5 (nL + 1) is 1.75g.
REQUIREMENT LESSONS LEARNED
305
MIL-STD-1797A
APPENDIX A
5.2.7.2 Pitch axis control power in maneuvering flight-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.7.2)
Flight test techniques are discussed under 5.2.8.1 ...steady-state control force per g.... Operational flight
test should be required for final demonstration, in order to assure that any problems are uncovered before
introduction into service.
Since demonstration is at constant throttle setting for 1-g equilibrium, these turns will involve loss of either
speed or altitude. While in a real sense sustained turn performance is a flying quality, its specification has
traditionally been a performance matter. The flying qualities requirement is on steady turn capability, with
throttle setting for 1-g flight, in order to avoid trim-change effects which are specified separately. Another
term sometimes used is instantaneous turn capability; properly that refers to a dynamic overshoot
capability.
VERIFICATION GUIDANCE
Critical flight conditions are indicated by the expression
Cm (α, q, δ, M, q , c.g.) + (zT / c ) [T/( q S)] = 0
where T is the 1-g trim thrust (which normally varies little with nz) and δ is a generic pitch control
deflection. Throttle setting is constant.
Measuring α to the principal axis and nz normal to the flight path.
δ=
− C L1
C mδ −
C mα
CLα
C Lδ
 Cm
I − I xp g c
C


α (n − 1) +  mq − zp
⋅
sin α cos α 

z
 4µ

n zmc 2
V2
 C Lα


m

1 
 n z −

n
z 

He

cos( α + ε) ⋅ 1 − 1/ n 2z 
mV c

where CL1 = W/( q S), µ = (W/S)/(gρ c ), sub p indicates principal axes, He is the engine angular
momentum Ieωe, nz =cosγ/cosφ) and ε is the inclination of the engine rotor axis to the principal x axis. In
the last term, the - is for right turns and the + is for left turns.
Initial flight safety restrictions to 80 percent of limit load may well dictate that the required Level 1 and 2
capability first be demonstrated in the course of the flight loads demonstration program. But the often
critical effects of angle of attack can still be investigated earlier, at higher altitude.
VERIFICATION LESSONS LEARNED
306
MIL-STD-1797A
APPENDIX A
4.2.7.2.1 Load factor response.
The time required to change from one level of normal load factor to another, in pullups and in turning
flight, shall be adequate for all maneuvers appropriate to the Flight Phase, for all conditions within the
Service Flight Envelope. Overshoots that result from abrupt pullups into the lift- or control-system-limited
region of the load factor boundary shall not result in departure or exceedance of load fact or limits.
REQUIREMENT RATIONALE (4.2.7.2.1)
Stating the operational load factor boundaries is not enough to ensure satisfactory maneuvering
response. Recent experience with F-5, F-15, and F-16 has shown that for aircraft that are departure
resistant, abrupt pitch commands for point and shoot capability or rapid deceleration provide increased
mission effectiveness. Use of relaxed static stability can require angle-of-attack/load factor limiters, pitchrate or angle-of-attack-rate anticipation features to be incorporated into the flight control system to ensure
that the limits are not exceeded. The effects of these features must be evaluated to ensure that loadfactor response is adequate.
REQUIREMENT GUIDANCE
There is no comparable MIL-F-8785C requirement.
There has been little work done to define satisfactory load-factor onset rates. The following, derived from
F-15 flight test data, can be used as a basis for comparison. The F-15 is considered to have good
response characteristics.
LOAD FACTOR RESPONSE
ALTITUDE
SPEED
RANGE
KNOTS EAS
PILOT CONTROL INPUT FROM
TRIMMED INITIAL CONDITIONS
RESPONSE
20,000
V=250
Step pitch control input from 1.5g turn
to max
Integral under body axis pitch
rate curve equals a change in
attitude of 32° in 1.1 sec
20,000
300<V<400*
Ramp pitch control input (approx.
10% max control surface rate) from a
trimmed 2g turn
From 2g to 7g in 1.8 sec
*NOTE: Between 300 and 400 kts, flight test data on an aircraft with a "g” limiter indicated that it took 3
seconds to go from 2 g's to 6 g's with a step input. The report indicated that this was inadequate for air-toground operations. All load factors are measured at the c.g.
REQUIREMENT LESSONS LEARNED
307
MIL-STD-1797A
APPENDIX A
5.2.7.2.1 Load factor response--verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.7.2.1)
Operational flight tests should be required for final demonstration, in order to assure that any problems
are uncovered before introduction into service.
VERIFICATION GUIDANCE
The critical flight conditions are dependent upon the characteristics of the particular vehicle.
VERIFICATION LESSONS LEARNED
308
MIL-STD-1797A
APPENDIX A
4.2.7.3 Pitch axis control power in takeoff.
The effectiveness of the pitch control shall not restrict the takeoff performance of the aircraft and shall be
sufficient to prevent overrotation during all types of takeoff . It shall be possible to obtain and maintain the
following attitudes during the takeoff roll: For catapult takeoffs ___________.
REQUIREMENT RATIONALE (4.2.7.3)
This requirement is intended to regulate against aircraft that exhibit no apparent pitch response to
commands during the takeoff roll until flying speed is reached (Vmin). These aircraft give no assurance that
rotation will be forthcoming, but then tend to “pop off", resulting in overrotation and a necessity for
immediate control reversal to avoid stall.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraph is 3.2.3.3.
The recommended attitudes are:
For nosewheel aircraft it shall be possible to obtain at 0.9 Vmin the pitch attitude that will result in a
lift-off at Vmin.
For tailwheel aircraft it shall be possible to maintain any pitch attitude up to that for a level thrust
line at 0.5 VS for Class I aircraft and at VS for Classes II, III, and IV.
These requirements shall be met on hard-surface runways. In the event that the aircraft has a mission
requirement for operation from unprepared fields, these requirements are to be met on such fields also.
This requirement is based on operational experience, which has shown that the ability to control pitching
attitude to achieve the proper attitude for lift-off before Vmin is necessary for acceptable liftoff at Vmin.
For catapult takeoff the recommendation is either not applicable (N/A) or
the effectiveness of the pitch control shall be sufficient to prevent the aircraft from
pitching up or down to undesirable attitudes in catapult takeoffs at speeds ranging from
the minimum safe launching speed to a launching speed 30 knots higher than the
minimum.
A related requirement in 4.2.9.1 is intended to provide adequate control of any overrotation tendency
without complicated control manipulation.
The requirement on takeoff control power is more important for single-engine aircraft or any multi-engine
aircraft which have VMCA (the airborne minimum control speed for a propulsion failure) equal to or less
than Vmin - For multi-engine aircraft where VMCA > Vmin, the requirement could be relaxed to 0.9 VMCA.
All aircraft will have to operate from hard-surface runways, and therefore hard surfaces were used as the
basic requirement. An increased coefficient of friction, however, such as occurs with unprepared fields,
increases the elevator effectiveness required for nosewheel aircraft but decreases the effectiveness
required of tailwheel aircraft, as can be seen in figure 96.
309
MIL-STD-1797A
APPENDIX A
FIGURE 96. Nosewheel and tailwheel lift-off.
The increased rolling friction force gives a nose-down pitching moment about the aircraft c.g. Nose-wheel
lift-off speed will increase monotonically with increasing µ, approaching the speed for takeoff in the
ground attitude. But tailwheel lift-off speed will decrease the increasing µ until just the application of
takeoff thrust will rotate the aircraft at zero speed. Then a different technique would be required. The
value 0.9 Vmin is a compromise between early enough indication of controllability and minimization of any
tendency to overrotate.
The requirement for control in catapult takeoffs could also be applied to ski-jump takeoffs or use of a
jump-strut.
REQUIREMENT LESSONS LEARNED
Single-engine propeller-driven airplanes with a T-tail have been deficient in terms of nosewheel rotation
prior to lift-off. As a result, pilot acceptance is very poor. Takeoff performance over an obstacle has been
demonstrated to be considerably worse in one T-tail aircraft than in an identical aircraft with a
conventional horizontal tail. Delayed lift-off has been attributed to inability to rotate to the takeoff attitude
prior to Vmin . The root cause is that the horizontal tail is out of the propeller wake. Multi-engine aircraft
which are not normally lifted off until VMCA (which is usually above Vmin) do not have as strong a
requirement for nose rotation at 0.9 Vmin. As an indication, multi-engine airplanes with T-tails have
generally been found to be acceptable to pilots.
310
MIL-STD-1797A
APPENDIX A
On the F-18, the pitch control-deflection range of ±20 deg. moved with pitch trim position, but could never
exceed 24 deg. trailing-edge-up (-) stabilizer. It was possible to mis-set trim for takeoff as that -24 deg.
deflection could not be obtained when needed during takeoff. The pilot should always have full deflection
available.
The requirement is especially important for turbojet aircraft for which relatively large pitch attitudes are
required for lift-off.
5.2.7.3 Pitch axis control power in takeoff-verification.
Verification shall be by analysis, simulation and flight test.
VERIFICATION RATIONALE (5.2.7.3)
Normally the critical test will be to start the rotation, i.e., to at least balance the aircraft with the nosewheel
or tailwheel just touching the runway but bearing no load. The main gear extension at this condition,
which determines the attitude there, can be found from the lift and load-stroke equations. Control to
maintain the takeoff attitude should also be checked, for compliance with the limitation of 4.2.9.1 on
control travel during takeoff. In flight test, takeoff performance tests are monitored for flying qualities.
VERIFICATION GUIDANCE
The ability to comply with this requirement should be obvious during operational flight test. Special
emphasis should be placed on short-field takeoffs at the maximum forward and aft center-of-gravity limits
for lift-off and overrotation, respectively.
From the equation
q S c Cm TzT – (W- q SCL - Tsinξ) (x+µy) = 0
the lift-off speed is given by
1
2
ρ VLO
2


T 
zT / c
 sin ξ +

1 −
W
X / c + µY / c  
W 
= q LO =

Cm
S 
CL +


X / c + µY / c


where ξ = iT + α; Cm and CL are the aerodynamic pitching moment and lift coefficients, in ground effect, of
the entire aircraft with takeoff thrust and fully deflected pitch controller, X and Y are the horizontal and
vertical distances between the c.g. and the main-gear axle; and µ is the main-gear coefficient of rolling
friction.
VERIFICATION LESSONS LEARNED
A balance must be struck between adequate control and overrotation tendency, e.g. the DC-9-80 takeoff
performance tests. The test pilot should be alert to avoid tail-strike.
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4.2.7.4 Pitch axis control power in landing.
The pitch control shall be sufficiently effective in the landing Flight Phase in close proximity to the ground
that ____________.
REQUIREMENT RATIONALE (4.2.7.4)
This requirement insures that the aircraft can be pitched up sufficiently, in ground effect, to achieve the
guaranteed minimum landing speed. It also insures that the nosewheel or tailwheel can be gently lowered
to the ground during landing rollout.
REQUIREMENT GUIDANCE
The related MIL-F-8785C requirement is 3.2.3.4.
It is recommended that the following requirements be placed on pitch control: during landing flare and
rollout with the aircraft trimmed for the minimum recommended approach speed not to exceed 1.3 VS(L):
a. The geometry-limited touchdown attitude can be achieved at touchdown, or alternatively
b. The lower of VS(L) or the guaranteed minimum landing speed [Vmin(L)] can be achieved when
flaring from shallow (γ = -3 deg) and steep (γ = -6 deg) approaches and
c. The nosewheel can be gently lowered to the ground at speeds down to 0.9 Vmin(L) or
d. For tailwheel aircraft, the tailwheel can be gently lowered to the ground at 0.5 Vmin(L) for Class
I and 0.75 Vmin(L) for Classes II, III, and IV.
This requirement is to assure adequate pitch control during flare and rollout in ground effect. Elevator
effectiveness can be severely degraded in ground effect due to a decrease in downwash caused by
presence of the ground plane.
A shallow approach is specified to eliminate the possibility of performing most of the flare out of ground
effect. Steep approaches are also required, as they tend to result in firmer touchdowns, making it difficult
to keep the nosewheel from slamming to the ground (at forward c.g.). The -6 deg glide path is generally
attainable, but could be moderated where the mission usage warrants.
REQUIREMENT LESSONS LEARNED
The Mitsubishi MU-2 (a twin-engine turboprop) is well known for a rapid pitchover immediately at
touchdown. Service difficulties with flight instruments and avionics are felt to result from this high shock
environment.
The T-46 has no guaranteed landing speed; since its geometry-limited touchdown attitude is greater than
its, stall angle of attack, VS (L), governs. For vehicles augmented to counter degraded static stability, the
change in center of rotation from c.g. to main gear at touchdown might result in an uncontrollable situation
even if ample control power exists. The original Space Shuttle in-flight control configuration did not
provide a stable control loop or fast enough trim capability after main gear touchdown due to the change
in, control geometry. It was difficult to lower the nose wheel gently. Control system parameter switching at
touchdown was required to correct this. For the Shuttle, then, control effectiveness was more than just
moment-producing capability. Some augmented aircraft, the X-29 for one, put up with a great deal of
control surface movement while taxiing in order to assure safe transition between ground and air. In any
case the stability and control augmentation must function effectively and safely during bounces, hard
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landings, etc., though the aircraft response characteristics are quite different on the ground than when
airborne. A canard control surface must not be allowed to stall: a gust at touchdown would be hard on the
nose gear.
5.2.7.4 Pitch axis control power in landing-verification.
Verification shall be by analysis, simulation, and flight test.
VERIFICATION RATIONALE (5.2.7.4)
The requirement is now structured so that verification may largely, if not entirely, be a part of maximumperformance landing tests.
VERIFICATION GUIDANCE
No particular guidance is deemed required except to note that the maximum forward c.g. (generally
regardless of weight) defines the critical flight condition for nose-up control of a stable airframe.
Crosswind may reduce ground effect; flight test should be in calm air. In general, ground effects often do
not correlate too well between wind tunnel and flight. See 4.2.8.6.2 guidance for discussion of the related
control force requirement.
VERIFICATION LESSONS LEARNED
The MIL-F-8785B/C requirement called for obtaining VS(L), the guaranteed landing speed or the
geometry-limited touchdown attitude in close proximity to the ground. By dispelling the inference of
steady, level flight, the new recommended requirement is in some respects safer to demonstrate and
more closely attuned to operational practice.
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MIL-STD-1797A
APPENDIX A
4.2.8 Pitch axis control forces
This section contains the control force gradients and limits to be applied to the pitch controller. As a word
of introduction, several points must be made which are applicable to all the following force requirements:
a. In general, the force requirements of MIL-F-8785C are unchanged. This is due in part to lack
of good, valid test data to justify changes, even if anecdotal information should suggest such
a change is warranted.
b. The requirements should be considered to be stringently applicable for male pilots only.
There is almost no available data for setting requirements for female pilots; a limited amount
of data, reviewed below, suggests considerably lower limits would be needed. This of course
presents a dilemma in setting limits for aircraft expected to be routinely operated by both
male and female pilots.
c.
Maximum forces specified appear in most cases to be quite large for the weaker male pilots
for continuous operation.
d. Effects of stick (or wheel) geometry and position on maximum force capabilities are not
explicitly covered in any of the requirements, though it is obvious that control location will
affect maximum attainable forces. This can be seen in the discussion that follows.
In a review of past research, Lockenour (AFFDL-TR-78-171) discussed the effect of stick location on push
and pull capability (figure 97), and the effect of upper arm angle on push and pull strength for the 5th and
95th percentile male (figure 98). The data shown in these figures are from Human Engineering Guide to
Equipment Design for male Air Force personnel in the sitting position. As described in AFFDL-TR-78-171
these data show that:
... one's maximum force capability is not symmetric left and right and varies by about a
factor of two for forward and aft stick positions ... pull and push strength differ significantly
and ... the 5th and 95th percentile male strengths differ by as much as a factor of three....
Certainly a given stick force at the grip will feel heavier to the pilot for aft stick positions.
Also one must be very careful in correlating the acceptability of stick forces for various
aircraft to include the effect of stick location and maximum stick deflection. For instance,
the F-5A stick deflection is greater than that of the A-7D by more than a factor of 2. This
places the stick in a different location in the cockpit for maximum deflection.
A more recent study, AFAMRL-TR-81-39, presents a comparison between male and female strength
characteristics for operating an aircraft control stick. Table XVI summarizes the percentiles for maximum
forces exerted by 61 men and 61 women on an aircraft control stick during a 4 second static exertion with
the right hand only. The 5th percentile values of men and women from these tests are also shown in
figure 97 for comparison. It can be seen that the 95th percentile woman has approximately the same
performance as the 5th percentile man. As AFAMRL-TR-81-39 observes, the force limits in this handbook
may not be consistent with the capabilities of pilots.
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MIL-STD-1797A
APPENDIX A
FIGURE 97. Effect of arm/stick geometry on maximum push and pull capability by the right arm
for the 5th percentile male (Human Engineering Guide to Equipment Design).
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MIL-STD-1797A
APPENDIX A
FIGURE 98. Effect of upper arm angle on pull and push strength for the 5th and 95th
percentile male (Human Engineering Guide to Equipment Design).
TABLE XVI. Maximum forces exerted on aircraft control stick (lb) by men and women
(AFANIRL-TR-81-39).
CONTROL
STICK
DIRECTION
Stick Forward
(Push)
Stick Back
(Pull)
MEN
WOMEN
5TH
PERCENTILE
50TH
95TH
5TH
PERCENTILE
50TH
93
123
165
46
87
109
64
85
106
48
52
64
95TH
Figure 99 shows data from Human Engineering Guide to Equipment Design illustrating the effect of wheel
angle on maximum push and pull capability for the 5th percentile male. The data are again for male Air
Force personnel, using the right arm only; the wheel grips are 18 inches above the Seat Reference Point
(SRP) and 15 inches apart. Figure 99a shows the various wheel angles and positions from the SRP. The
greatest push and pull capability occurs at the furthest position of the wheel where the pilots' entire arm is
used. This can be seen in figure 99b where the push capability at 23-1/4 inches from the SRP is
approximately twice that obtained when the control wheel is at its closest at 10-3/4 inches from the SRP.
Similarly the maximum pull capability varies almost by a factor of 2 in figure 99c depending on the control
wheel angle and position.
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MIL-STD-1797A
APPENDIX A
FIGURE 99. Effect of arm position and wheel angle on maximum push and pull capability by the
right arm for the 5th percentile male (Human Engineering Guide to Equipment
Design).
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MIL-STD-1797A
APPENDIX A
It must be stressed that these are maximum forces in single applications; clearly, continuous operation
(such as would be expected in meeting any of the force requirements) would produce much lower
maximum forces. In a discussion of some general principles of control design, and one- vs. two-handed
operations, Human Engineering Guide to Equipment Design states that:
For controls requiring single applications of force, or short periods of continuous force, a
reasonable maximum resistance is half of the operator's greatest strength. For controls
operated continuously, or for long periods, resistances should be much lower... Controls
requiring large forces should be operated with two hands (which, for most controls, about
doubles the amount of force that can be applied) depending on the control type and
location and on the kind and direction of movement as follows:
When two hands are used on a stick ... located along the body midline, pull is generally
almost doubled. Push is doubled near the body but is only slightly stronger at distances
away from the body...
When two hands are used on stick ... controls located on either side of the body midline,
at or beyond the shoulder, pull is approximately doubled, push is not greatly increased
except at close distances...
Since we expect to use this standard for few aircraft with unpowered controls, little design penalty would
result from lowering the maximum allowable forces. Stability and control augmentation, and response feel
systems, could still keep the lightest force gradients within the presently allowable limits. Concern has
been expressed, however, that with lighter maximum forces some heavy-handed pilots might be inclined
to overstress the vehicle. In particular that might be the fate of trainers with new student pilots. Then,
especially in a simple aircraft, a lower gradient at forward c.g. might make the gradient too low at aft c.g.
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MIL-STD-1797A
APPENDIX A
4.2.8.1 Pitch axis control forces--steady-state control force per g.
These requirements apply for all local gradients throughout the range of Service load factors defined in
4.1.4.2. The term gradient does not include that portion of the force versus n curve within the breakout
force.
a. In steady turning flight and in pullups and pushovers at constant speed, the incremental
control force required to maintain a change in normal load factor and pitch rate shall be in the same
sense (aft -- more positive, forward -- more negative) as that required to initiate the change.
b. The variations in pitch controller force with steady-state normal acceleration shall have no
objectionable nonlinearities within the following load factor range: _____. Outside this range, a departure
from linearity resulting in a local gradient which differs from the average gradient for the maneuver by
more than 50 percent is considered excessive, except that larger increases in force gradient are
permissible at load factors greater than 0.85nL.
c. The local force gradients shall be: __________. In addition, FS/n should be near the Level 1
upper boundaries of these gradients for combinations of high frequency and low damping.
For all types of controllers, the control force gradients shall produce suitable flying qualities.
REQUIREMENT RATIONALE (4.2.8.1)
These requirements relate to the classical stick-free maneuvering stability (stick force per g, FS/n) at
constant speed. The basic premise is that FS/n represents a necessary tactile cue for pilot regulation of
load factor. Low values of FS/n result in excessive sensitivity with a tendency toward exceeding the
aircraft structural limits. High values lead to pilot fatigue during maneuvering.
REQUIREMENT GUIDANCE
The related MIL-F-8785C paragraphs are 3.2.2.2 and 3.2.2.2.1.
Recommended ranges for limiting departures from linearity are:
CLASS
MIN. nz
MAX. nz
I, II & III
0.5
Lesser of
0.5[no (+) + 1] or 3
IV
0
3
The requirements for control forces in maneuvering flight, unchanged from MIL-F-8785C, are listed in
table XVII.
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MIL-STD-1797A
APPENDIX A
TABLE XVII. Pitch maneuvering force gradient limits.
Level
Maximum Gradient
(Fs/n) max, lb/g
Minimum Gradient
(Fs/n) min, lb/g
1
240/(n/α)
but not more than 28.0
nor less than 56/(nL – 1) *
The greater of
21/(nL – 1)
and 3.0
2
360/(n/α)
but not more than 42.5
nor less than 85/(nL – 1) *
The greater of
18/(nL – 1)
and 3.0
3
56.0
The greater of
12/(nL – 1)
and 2.0
a. Center stick controllers
* For nL <3, (fs/n)max is 28.0 for Level 1, 42.5 for Level 2.
b. Wheel controllers
1
500/(n/α)
but not more than 120.0
nor less than 120/(nL – 1) *
The greater of
35/(nL – 1)
and 6.0
2
775/(n/α)
but not more than 182.0
nor less than 182/(nL – 1) *
The greater of
30/(nL – 1)
and 6.0
3
240.0
5.0
It was decided that the major differences in the desired maneuvering forces between fighter aircraft and
transports are due to the type of controller, in addition to aircraft Class. The effects of aircraft class (really
a grouping of types of missions) seem to be adequately described by limit load factor, through the K/(nL 1) formulas of MIL-F-8785C. In addition, however, there are several arguments for having different
maneuvering forces for centerstick and wheel controllers. For example, the lower limits on maneuvering
forces must be higher with a wheel control because the pilot's arm is usually unsupported, whereas the
pilot has very good vernier control with a centerstick even with light forces because his forearm is partially
supported on his thigh. In any case, pilots seem to agree that they cannot maintain the precision of
control with a wheel that they can with a stick, and that the maneuvering control forces should be higher
for the wheel.
There is some evidence (ASD-TDR-63-399 and AFFDL-TR-67-51) that FS/n at very low n/α can or should
be higher than at high n/α. This is possibly due to a gradual change from concern with load factor and
structural protection at high n/α to concern with control of pitch attitude alone at low n/α. Specification of
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MIL-STD-1797A
APPENDIX A
forces in the form of limits on FS/α at low n/α can be accomplished by making the FS/n limits vary
inversely with n/α: at constant speed,
FS/α = (FS/n) (n/α)
On the basis of these considerations, the upper limits on Fs/n were expressed in the form K/(nα) at low
n/α and K/(nL -1) at high n/α, with separate requirements for stick and wheel controllers. On the basis of
long experience with unpowered-control airplanes, which tend to have Fs/ invariant with airspeed, the
lower boundaries do not vary with n/α.
However, there is some question as to the significance of the ASD-TDR-63-399 and AFFDL-TR-67-51
tests. These references are discussed in detail under Supporting Data.
To illustrate the use of the gradient limits of table XVII, figures 100 and 101 show possible boundaries for
two representative aircraft. Figure 100 is for a centerstick controller with nL = 7.0; figure 101 is for a wheel
controller with nL = 3.0. Similar plots may be constructed for any aircraft using the table XVII formulas.
However, such plots, while representing the table XVII suggested limits, do not convey the entire picture,
as illustrated by the following considerations.
Effects of stick/wheel position on acceptability of FS/n are not covered by these requirements. But with the
variation in force capability shown in the discussion of 4.2.8 it seems intuitively obvious that there must be
an interrelationship of force and deflection gradients with controller location.
Another item for consideration is the allowance in table XVII for considerably higher values of FS/n when
n/α is low. For example, figure 102 (from AFFDL-TR-78-171) illustrates the variation of FS/n with airspeed
for the OV-10A aircraft, due to an elevator spring tab. The apparently large and rapid change in FS/n
actually results in fairly constant maximum stick force at stall (figure 103) or nL, whichever comes first.
Perhaps that becomes the important factor at speeds below maneuvering speed (VA, see figure 102).
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MIL-STD-1797A
APPENDIX A
FIGURE 100. Elevator maneuvering force gradient limits: center-stick controller, nL = 7.0.
FIGURE 101. Elevator maneuvering force gradient limits: wheel controller, nL = 3.0.
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MIL-STD-1797A
APPENDIX A
FIGURE 102. OV-10A maneuvering control (AFFDL-TR-78-171).
FIGURE 103. Longitudinal stick force at stall (AFFDL-TR-78-171).
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MIL-STD-1797A
APPENDIX A
A more basic consideration relative to table XVII is the complete absence of a force gradient specification
for sidestick controllers, reflecting in part the limited data base. However, AFFDL-TR-79-3126 and Class
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MIL-STD-1797A
APPENDIX A
79A Final Report of AF Test Pilot School, based on a series of flight tests conducted by students of the
USAF Test Pilot School, give some insight into preferred gradients. These data, figure 104, are for an airto-air tracking task. The test aircraft was the USAF/Calspan variable stability T-33, with a T-38A utilized
as the target. The ratings shown are the average over three pilots, all with fighter experience. An
approximate Level 1 boundary is suggested in figure 104. In general, the Fs/n range is comparable to that
of table XVII, i.e., 2-14 lb/g. The relatively weak frequency dependence may also exist for centerstick
controllers, although there are no data to support this. More data from these evaluations are given in
Requirement Guidance for 4.2.8.4.
FIGURE 104. Short-period frequency vs longitudinal stick force per g (FS /δ
separately optimized) for a side-stick controller.
The FS/n gradients in figure 104 are the initial values; as mechanized, the slope at larger deflections was
half the initial slope. The breakout forces were 1/2 lb. AFFDL-TR-79-3126 suggests that the sidestick
neutral position
… be oriented so that in wings-level unaccelerated flight the pilot need never move his
wrist further aft than 5-7° forward of vertical to command maximum permissible load
factor...
Available data would tend to support a neutral position of 10° to 17° forward of vertical
and 8° to 12° left (inboard) of vertical.... A pilot adjustable armrest is absolutely
mandatory, and its design can influence pilot acceptability as much as any other
parameter.
SUPPORTING DATA
The data base for control force requirements is sparse, and limited entirely to Class IV aircraft. The most
thorough data sets are from USAF/Calspan T-33 flight tests in which pilots chose optimum values of FS/n
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MIL-STD-1797A
APPENDIX A
for varying short-period characteristics. However, the useful information is basically a byproduct, since the
intent of these tests was the study of short-period frequency and damping. So, while specific conditions
can found for which ζsp and ωsp fell w